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Introduction to Space Science and Spacecraft Applications

Introduction to Space Science and Spacecraft Applications

INTRODUCTION TO SPACESCIENCES AND SPACECRAFT APPLICATIONS

INTRODUCTION TO SPACESCIENCES AND SPACECRAFT APPLICATIONS

BRUCEA. CAMPBELL AND SAMUEL WALTER MCCANDLESS,JR.

Gulf Publishing Company Houston, Texas Introduction to Space Sciences and Spacecraft Applications

Copyright 0 1996 by Gulf Publishing Company, Houston, Texas. All rights reserved. Printed in the United States of America. This book, or parts thereof, may not be reproduced in any form without permission of the publisher.

Gulf Publishing Company Book Division P.O. Box 2608 Houston, Texas 77252-2608

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Library of Congress Cataloging-in-Publication Data Campbell, Bruce A., 1955- Introduction to space sciences and spacecraft applications / Bruce A. Campbell, Samuel Walter McCandless, Jr. p. cm. Includes bibliographical references and index. ISBN 0-88415-411-4 1. Astronautics. 2. Space vehicles. I. McCandless, Samuel Wal- ter. 11. Title. TL791.C36 1996 629.4-dc20 96- 14936 CIP

Printed on acid-free paper (=)

iv Contents

Acknowledgments viii About the Authors viii Preface ix

CHAPTER 1 Introduction and History 1 Uses of Space, 1. , 3. ReferencedAdditional Reading, 23. Exercises, 23.

PART 1 SPACE SCIENCES

CHAPTER 2 Orbital Principles 26 Orbital Principles, 28. Orbital Elements, 38. Orbital Properties, 39. Useful , 44. Establishment and Orbital Maneuvers, 47. ReferencedAdditional Reading, 5 1. Exercises, 5 1.

CHAPTER 3 Propulsion 53 First Principles, 54. Orbit Establishment and Orbital Maneuvers, 63. ReferencedAdditional Reading, 74. Exercises, 74.

CHAPTER 4 Spacecraft Environment 76 The , 76. The , 87. Spacecraft Effects, 97. ReferencedAdditional Reading, 101. Exercises, 101.

V PART 2 SPACECRAFT APPLICATIONS

CHAPTER 5 Communications 104 Communications Theory, 104. Modulation, 115. Digital Communications, 118. Communications Systems, 128. ReferencedAdditional Reading, 129. Exercises, 129.

CHAPTER 6 Remote Sensing 132 Remote Sensing Principles, 135. Measurement Dimensions of Remote Sensors, 146. Remote Sensing , 149. ReferencedAdditional Reading, 150. Exercises, 151.

CHAPTER 7 Navigation 154 Position Determination Using Doppler Techniques, 154. Pulse Ranging and Phase Difference Positioning, 159. ReferencedAdditional Reading, 164. Exercises, 164.

PART 3 SPACECRAFT SYSTEMS AND DESIGN

CHAPTER 8 Spacecraft Systems 166 Satellite Design, 166. Ground Support Systems, 192. Space System Evolution, 195. ReferencedAdditional Reading, 196. Exercises, 197.

CHAPTER 9 Spacecraft Design 198 The Systems Approach, 198. Spacecraft Systems, 209. ReferencedAdditional Reading, 210. Exercises, 2 10.

vi APPENDIX A Manned Spaceflight Summary 21 1

APPENDIX B United States and Foreign Spacecraft Inventories and Spacecraft Descriptions 21 7 Index 230

vii Acknowledgments

The authors express appreciation to all the colleagues and students whose reviews and comments helped to shape and refine this book.

About the Authors

Bruce A. Campbell is currently involved in spacecraft project man- agement at NASA's Goddard Space Flight Center. He is an aerospace engineering graduate of the U.S. Naval Academy and flew carrier-based jet aircraft as a Naval Flight Officer. Samuel Walter McCandless, Jr., is president and founder of User Sys- tems, Inc., which provides airborne and space-based radar and microwave sensor and system design and evaluation, remote sensing and satellite sys- tem design, and associated services. He was test director for NASA's Sur- veyor 1 spacecraft and managed NASA's Seasat program from its incep- tion through operation of the satellite in space.

viii Preface

In 1985, the authors were asked to create an introductory course for a new series of classes focusing on space systems engineering in an under- graduate aerospace engineering curriculum. Although the authors were familiar with separate texts on each of the subjects to be taught, there seemed to be no single existing text that covered all the topics at a level that could be used for an introductory class. The first time the course was taught, handouts were created for each of the topics covered. Each suc- cessive time that the course was taught, the handouts were updated based on the experience of the preceding class. The idea to publish came very early, and the handouts evolved into the chapters that make up this text. In its manuscript form, this text has been used by several hundred stu- dents, in several different institutions, under different instructors, with very favorable reviews. The most common comment is the ease with which the student can read and understand the material. Although the text is geared toward sophomore/junior undergraduate engineering students who will continue to take courses in these space-related topics, it has proven to be comprehensible and interesting to students in other science and nonengineering fields as well. The subjects are introduced on a basic level with no prior related knowledge expected of the student. A funda- mental knowledge of physics, differential equations, dynamics, and other pre-engineering subjects is helpful, but not necessary, to understanding the basic concepts presented. To emphasize this point, the subject matter of the text has been con- densed into a “short course” that has been presented to many diversified groups of managers, technicians, military personnel, and other profes- sionals-not necessarily engineers-over the past several years. Many of these people are involved in space systems acquisition or operation, and most report that the course provides them with a more complete level of understanding that makes them more comfortable in their fields and which they believe will help them in their professions. The text material has proven to be an excellent reference and review source for the student and professional alike. This text considers many of the interdisciplinary topics necessary for understanding the design and application of space-based systems. The

ix space science topics include orbital mechanics (Chapter 2), propulsion (Chapter 3), and a description of the space environment (Chapter 4). These chapters prepare the way for discussing spacecraft applications. The most common uses of systems in space include communications (Chapter 5), remote sensing (Chapter 6), and navigation (Chapter 7). Finally, the basic systems required by most spacecraft (Chapter 8) and a methodology used to design a spacecraft (Chapter 9) are presented. In a four-year, semester-oriented undergraduate institution, it is recom- mended that this text be used for a 3-credit-hour (2 hours classroom, 2 hours laboratory) course given at the end of the sophomore year or begin- ning of the junior year. By this time the student should be familiar with most of the helpful subjects discussed above and will be ready to take this course and, if desired, any of the follow-on space topic courses. Early in the course, the laboratory hours can be used for additional classroom hours and problem solving; however, as the course progresses, laboratory hours could include orbital mechanics computer simulations, Estes rock- et firings, an “outside” environmental lab, a communications lab, review of some remote sensing materials or films, a GPS navigation demonstra- tion, or other demonstrations that help the student get a feel for the sub- jects discussed. Later in the course, the laboratory hours should be devot- ed to the formulation of an idea for a space-based application and the top-level conceptual design of a system to fit the need. This text could also be used for a 2- or 3-hour classroom-only course which might be more suitable to the nonengineering student or two-year institution.

X CHAPTER 1 Introduction and History

To set the stage for what is presented in this text, a brief discussion on the uses of systems in space is provided as insight into the operation and purposes of such systems. This discussion is followed by a brief history of man’s achievements in space and some conjecture on where we may be headed.

USES OF SPACE

For centuries, the stars and planets and their patterns and positions have been used for purposes similar to those for which we use modem space sys- tems. Notable and ancient applications include navigation and environmen- tal prediction. So it is surprising that even after Sir Isaac Newton showed that man could, theoretically, place an object into space, no one immediate- ly thought of any useful reason why anybody would want to. It was not until this feat was actually accomplished that we began to explore and discover the benefits offered by presence in space and began designing and building systems to take advantage of this wonderful new dimension. After the first few pictures were returned from some of the early sub- orbital test rocket flights, experimenters recognized that one of the great- est advantages of being in space was the expanded perspective looking back toward earth. The ability to view areas of the globe large enough to show entire storm patterns spawned the science of global environmental remote sensing. This has dramatically increased knowledge of the envi- ronment, allowing for the observance and prediction of adverse condi- tions and thereby saving life and property from disaster. Remote sensing systems also allow us to discover and monitor natural resources and eval- uate demographic effects such as urban expansion and pollution. Similar type sensors, positioned outside the blanket of the earth’s atmosphere and

1 2 Introduction to Space Sciences and Spacecraft Applications on interplanetary craft, allow a clearer look at the universe of which we are such a small part, giving us a better understanding of where we have been and what the future may hold. The need for advanced global communication capabilities caused many to look toward space as an alternative to earlier methods. As early as the 1940s, science writer Arthur C. Clarke hypothesized that a platform placed at a particular location in space (where it would appear motionless with respect to the ground) could be used as a communications relay sta- tion. A multitude of such systems exist today, and the resulting real-time, global availability of information has transformed the news, entertain- ment, and communications industries and has had a major impact on polit- ical and military operations as well. When the first artificial satellite, the ’s I, began circling the globe, U.S. scientists discovered a method for using the sig- nals transmitted by the satellite to accurately determine one’s position anywhere on the earth. Satellite navigation is now employed by military, government, private, and commercial users for positioning and surveying on land as well as for aircraft and ship navigation. When man himself ventured into space, many benefits resulted as by- products of the developments in technology required to support these mis- sions. conducted experiments that proved that the unique envi- ronment of space offered many useful possibilities. In microgravity

Figure 1-1. Sputnik (left) and Explorer (rightethe first satellites launched by the Soviet Union and the United States, respectively. Introduction and History 3 environments, materials processing impossible in the -constrained laboratories on earth is possible and may result in new alloys, crystals, and medicines which could improve man’s existence. The need for electrical power for space systems has lead to alternative sources of energy for use on earth. The future uses of space are still developing, and space may truly become mankind’s next frontier.

HISTORY OF SPACEFLIGHT

Even though many consider the launch of Sputnik I by the Soviet Union as heralding the Space Age, man’s efforts toward this endeavor actually began much earlier. Notwithstanding the dreams of writers such as H. G. Wells and Jules Verne, the first practical efforts toward spaceflight really began with early rocket pioneers such as Russia’s Konstantin Tsiolkovsky, Germany’s Hermann Oberth, and America’s Robert Goddard. The V-2 rocket, used to terrorize England during World War 11, was the first opera- tional rocket system. America’s advances began after the war with the aid of expatriated German scientists, including Wernher von Braun, and a cache of captured V-2s with which to experiment. The Soviet Union received the same type of reparations after the war and began a similar effort. The first efforts by both countries were in developing improvements to the V-2, and well before Sputnik I, both countries had operational inter- continental ballistic missile (ICBM) systems. Many of the V-2 rockets and follow-on booster designs carried scientific experimental packages on board during test flights. These packages measured the characteristics of the atmosphere, the ionosphere, and the low-altitude space environment. One such package was used by Dr. James Van Allen to discover the exis- tence of regions of highly energetic particles encircling the globe, subse- quently named the Van Allen radiation belts. The Soviet Union placed the first man-made object into orbit around the earth on October 4, 1957. The overhead signals received from the 83-kg (184-lb) Sputnik I (“Traveler”) alarmed many who demanded that the United States match the feat. Pressure increased when a month later Sput- nik 2, at a surprising 507 kgs (1,120 lbs), went into orbit carrying a small dog to test the effects of space on a living creature. (The dog Laika sur- vived for a week until purposely poisoned, because the craft was not capable of reentry.) America finally succeeded with Explorer I, launched on January 31, 1958, but U.S. capabilities were severely questioned when comparing the 8-kg (18-lb) satellite to the two prior Sputniks. 4 Introduction to Space Sciences and Spacecraft Applications

Figure 1-2. Launch of Alan Shepard, America's first man in space, on the Mer- cury-Redstone-3 mission, May 6,1961. (Photograph courtesy of NASA.)

In an effort to consolidate and accelerate America's efforts in space, Congress authorized the establishment of the National Aeronautics and Space Administration (NASA) in July 1958. The new agency absorbed the National Advisory Committee for Aeronautics (NACA) as well as many military space projects, including their personnel and facilities. Both countries continued to launch larger and more sophisticated Introduction and History 5 unmanned satellites for a variety of military and scientific purposes, but it was apparent that efforts were also being made to place a man into space.

Manned Spaceflight

Once again, the Soviets were first in the manned spaceflight effort with the launch of Air Force Major Yuri A. Gagarin in the 4,706-kg (10,400-lb) Vosfok 1 (“East”) spacecraft on April 12, 1961, for a single orbit. The recovery area, unlike the later U.S. manned flights, was back in the Soviet homeland, and, after the spacecraft reentered and was on its way down by parachute, Gagarin exited the capsule and used his own parachute to land! Three weeks later on May 5, 1961, the United States launched Navy Commander Alan B. Shepard, Jr., into space. The successful 15-minute suborbital flight reached an altitude of 187.6 km (116.5 mi) and went 489.1 km (303.8 mi) downrange to recovery in the South Atlantic. Although successful, this flight showed, once again, that America was lagging behind the Soviet Union in what was now being called the “Space Race.” In an effort to turn this tide, President John E Kennedy, during a major speech before Congress only three weeks later, publicly committed the United States to landing a man on the , and returning him safely to earth, before the decade was out. With this clear objective in mind, the U.S. manned space program took shape, evolving from the existing Mer- cury program, through Gemini, to Apollo.

Mercury. The purpose behind the Mercury program was to gain basic data on the effects of spaceflight on human beings. To gain preliminary physical data as well as test the launcher and spacecraft systems, a rhesus monkey (Sam, December 1959) and a chimpanzee (Ham, December 1960) were launched on suborbital flights, and another chimp (Enos, November 1961) was launched into orbit before these same flights were attempted by astronauts. Alan Shepard’s flight was followed by a second suborbital flight carry- ing Virgil “Gus” Grissom in July 1961. John Glenn became the first American to orbit the earth on February 20, 1962. Three additional orbital Mercury flights followed, the longest by Gordon Cooper who spent more than 34 hours alone in orbit.

Gemini. Gemini flights were designed to evaluate the ability of perform- ing the tasks in space required for a manned lunar landing. Three methods 6 Introduction to Space Sciences and Spacecraft Applications

Figure 1-3. Photograph of the two-man Gemini 7 spacecraft taken from the Gemini 6 spacecraft during rendezvous maneuvers in December 1965. (Pho- tograph courtesy of NASA.) of conducting the moon flights were under consideration. One was the direct ascent of a spacecraft to the lunar surface and back. Another was known as earth-orbital rendezvous, by which two spacecraft would be launched separately to rendezvous in the earth’s orbit before continuing as one to the moon and back. The method ultimately selected was called lunar-orbit rendezvous, by which a spacecraft would be launched into lunar orbit whereupon a specially designed lunar landing module would travel down to the surface and back up to lunar orbit to redock with the mother ship for the return to earth. In order to accomplish this, crew size would have to be increased, longer flight duration would be necessary, and in-orbit operations such as extravehicular activities (EVAs) and ren- dezvous and docking techniques would have to be perfected. The Gemini capsule was basically a modified Mercury capsule, enlarged to carry two men, with the capability to perform the required on- orbit operations. There were ten Gemini flights in the two years the pro- gram ran from 1965 to 1966. These flights were a significant evolution Introduction and History 7 from the earlier Mercury flights in which the astronauts were basically only along for the ride, and the outstanding successes of Gemini gave NASA confidence in its abilities to continue on toward the moon.

Apollo. Unlike the earlier flights which used modified ICBMs as launch- ers, Apollo required a totally new designed specifically to carry men and equipment to the moon. In 1963, even before the first Gem- ini flights, Wernher von Braun’s team at the Marshall Space Flight Cen- ter in Huntsville, Alabama, was given the go-ahead to develop his Saturn family of powerful boosters. The rocket still represents the most powerful launch vehicle ever developed in the United States and was a “giant leap” in technology at the time. However, the confidence gained from the Gemini flights was shaken in January 1967 when a fire broke out in the Apollo I capsule during a prelaunch test, killing astronauts Gus Grissom, Ed White, and Roger Chaf- fee. Redesign of the capsule to prevent any similar occurrences delayed the first manned Apollo fight, designated , until October 1968. Only nine months later, on July 20, 1969, astronauts Neil Armstrong and Edwin “Buzz” Aldrin entered the gangly-looking lunar excursion module (LEM) and descended to the surface of the moon while Mike Collins orbited overhead in the Apollo 11 command module. Between this date and December 1972, five more crews landed on the moon. The twelve astronauts deployed experiments, explored the surface, and returned a total of 378 kgs (836 lbs) of lunar material for study by the international scientific community. The only other mishap occurred on Apollo 13, when an oxygen tank in the service module section of the spacecraft exploded on the way to the moon. The tense situation had a happy ending when the crew was recov- ered six days after launch. having used the lunar module as a lifeboat and booster rocket to push themselves around the moon and back toward the earth in place of the ailing commandservice module. The Apollo program had successfully met its objectives, and the Unit- ed States had clearly accomplished an enormous feat. But what next?

Skylab. After the lunar landings, public and governmental support for the space program diminished. With the escalation of the Vietnam War, peo- ple began to question the relative worth of the 25 billion dollars spent for the program. Three planned lunar landings (Apollo 18, 19, and 20) were canceled as a result of a review of the nation’s space program by a Space 8 Introduction to Space Sciences and Spacecraft Applications

Figure 1-4. Apollo 15 on the surface of the moon with the lunar excursion module and lunar rover vehicles nearby. (Photograph courtesy of NASA.)

Task Group of the National Academy of Sciences established by President Richard M. Nixon in 1969. The new focus would be on development of a reusable space transportation system. Initial plans also called for two separate space stations to be visited by multiple crews using Apollo-derived equipment. Instead, a single station, Skylab, was approved. When it was launched atop the last Saturn V boost- er in May 1973, a meteoroid shield and one of the two main solar panels were lost during ascent. The first crew to arrive found living conditions unbearably hot and a station lacking power. Through ingenuity and unplanned EVAs, the crew was to erect a sunshield and straighten out the power problem to make the station usable. The longest mission to Sky- lab was by the third and final crew in 1974 and lasted 84 days. Skylab was used for a wide variety of investigations, including pio- neering earth and atmosphere remote sensing, materials processing, bio- medicine, and some of the most revealing solar studies to date. Aban- Introduction and History 9 doned after a total of 171 days in use, Skylab’s orbit gradually decayed until it finally reentered the atmosphere in July 1979 and was destroyed.

Apollo-. As an act of dCtente, the United States and the Soviet Union agreed to a joint space operation involving the docking of an Apol- lo capsule with a Soyuz spacecraft. Scientists, engineers, and crews from both countries worked together to design a universal docking adapter to enable the two ships to dock and give direct access between the space- craft. Both craft were launched on the same day in July 1975 and docked two days later. In addition to ceremonies and interviews for television, the crews conducted joint experiments using facilities on board both space- craft. Though not technologically critical, the Apollo-Soyuz mission demonstrated that the two superpowers could cooperate in the peaceful use of outer space.

Space Shuttle. From 1975 to 1981, no American manned missions were launched. NASA focused on the development of the reusable space trans- portation system (STS) commonly referred to as the . The

Figure 1-5. Three crews visited Sky/ab between May 1973 and February 1974. (Photograph courtesy of NASA.) 10 Introduction to Space Sciences and Spacecraft Applications

n

Figure 1-6. Apollo/Soyuz: The United States and Soviet spacecraft remained docked for two days while the crews performed joint experiments.

shuttle was seen as the next step in spaceflight evolution, allowing routine access to space. New satellite designs would take advantage of the large payload compartment and relatively easy trip into orbit compared to con- ventional launchers. Satellites and experiments could be deployed, ser- viced and repaired, retrieved, and returned to earth by the shuttle. The flight crew, which would include nonpilot scientists and technicians, could conduct experiments and observations in a shirt-sleeve environment or don space suits to conduct activities outside. In its original conception, the Space Shuttle was to take off and land at conventional runways and be completely reusable for up to 100 flights, at least halving the cost per pound of carrying payloads into space. Howev- er, design difficulties and budgetary constraints lead to a simpler and less expensive configuration using solid rocket motors and an expendable external fuel tank to boost an orbiter into space which, after its mission was complete, would glide back to earth to land like an aircraft. On April 12, 1981, the first space shuttle, Columbia, lifted off from Kennedy Space Center for a completely successful two-day mission with astronauts John Young and Robert Crippen on board. In a little under five years, the four-ship shuttle fleet flew 24 successful missions carrying satellites, experiments, and up to a seven-person crew into low-earth orbit. However, the twenty-fifth mission, launched on January 28, 1986,

12 Introduction to Space Sciences and Spacecraft Applications

,, / /

Figure 1-8. Only the Soyuz CEW compartment is movered after a flight.

Over two years elapsed before the next Soviet manned launch, which was the first flight of the new Soyuz (“Union”) spacecraft only three months after the ApoZZo I fire. This first flight ended in tragedy when the recovery parachute lines became tangled on reentry and the lone cosmo- naut was killed on impact. Like Apollo, the Soyuz program was set back for more than a year before the next manned flight. Evidence that the Soviets were working toward a manned landing on the moon came when a spacecraft, launched in 1971, reentered and crashed in Australia in 198 1. Concern over an earlier reentry of a radioac- tive nuclear power generator forced the Soviets to reveal that the mission had been a test of a lunar cabin. Lack of a launch vehicle with sufficient thrust seems to be the reason for the slow Soviet progress in this effort, which was then abandoned after the U.S. successes. From 1971 to 1982, six successful Salyut space stations were launched and visited by a total of 29 crews. About one-quarter the size of Skylub, the stations were used for a number of science, biomedical, and military missions, the longest lasting 237 days aboard Salyut 7 in 1984. The fol- lowing space station design, Mir (“Peace”) was a step forward with its Introduction and History 13 capability for longer missions and expansion and modification. Crews have remained on board for up to a year, and several docking and labora- tory modules have been added to the basic core. The Russians continue to use the Soyuz spacecraft (with modifications) to this day, concentrating on earth-orbit operations. For the past several years they have averaged two to three Soyuz flights per year, exclusively associated with visits to the Mir space station.

International Manned Spaceflight. The United States and Russia remain the only countries capable of launching and supporting manned spaceflight missions independently. However, individuals from several other countries have participated in various U.S. and SoviedRussian manned missions. The Space Shuttle has carried several international pay- loads accompanied by payload specialists and scientists from these host countries. Besides cosmonauts from the former Soviet Union states, the Intercosmos program carried “guest” cosmonauts from other countries into space aboard the Soyuz capsule. Countries that have participated in these programs include Afghanistan, Austria, Bulgaria, Canada, Cuba,

Ill I I Ill I I Ill I I Ill I I Ill I I Ill I I Ill I I I i B

Figure 1-9. The Mir space station has been in near-continuous use since its launch in February 1986. 14 Introduction to Space Sciences and Spacecraft Applications

Czechoslovakia, France, Germany, Hungary, India, Japan, Mexico, Mon- golia, the Netherlands, Poland, Romania, Saudi Arabia, Syria, the United Kingdom, Vietnam, and the United States. A listing of manned spaceflight missions, crews, and mission highlights through the first flight of the Space Shuttle is given in Appendix A.

Unmanned Spaceflight

In addition to manned spacecraft, the United States, the Soviet Union/ Russia, and, more recently, many other countries have developed, launched, and operated hundreds of unmanned spacecraft designed to per- form military, national interest, scientific, and commercial applications. The following sections highlight some of the application areas in which these spacecraft perform and the evolution of spacecraft within each area.

Earth Observation. Due to the unique vantage point of space, many satellites are designed to observe the earth to collect data on the environ- ment. TZROS I (Television Infrared Observation Satellite), the first weath- er satellite, was launched in April 1960 and returned over 22,500 pictures of the earth. TIROS, NOM (National Oceanographic and Atmospheric Administration), GOES (Geostationary Operational Environmental Satel- lite), and the military DMSP (Defense Meteorological Satellite Program) systems monitor the earth’s weather, allowing timely prediction of cli- matic changes and their effects on the planet. In September 1991, the Upper Atmosphere Research Satellite (UARS) was deployed from the Space Shuttle to conduct the first comprehensive studies of the stratosphere, mesosphere, and thermosphere regions of the atmosphere. The total uzune mapping spectrometer (TOMS), a NASA instrument that “hitched a ride” on U.S. Nimbus and SovietRussian Meteor weather satellites, discovered and continues to monitor the “ozone hole” region of dangerously depleted ozone levels which appears over Antarctica each fall. Missions such as these allow researchers to map the short- and long-term changes occumng in the atmosphere and to deter- mine what effects may be attributable to humans. ERTS I (Earth Resources Technology Satellite) was launched in July 1972 into an orbit over the poles and used high-resolution electronic cam- eras to collect imagery on geology, crops, population, and pollution. ERTS satellites evolved into the Landsat series which continues to focus Introduction and History 15

Figure 1-10. The Geostationary Operational Environmental Satellite (GOES). on discovering and monitoring earth’s resources in even the remotest areas of the world. In June 1978, Seasat was launched to gather information on sea tem- perature, sea ice, wind speed and direction, and wave heights. Though its mission was cut short due to an electrical failure, Seasat first demonstrat- ed the beneficial use of synthetic aperture radar (SAR) to provide all- weather, day and night, high-resolution images for a variety of earth-mon- itoring applications. A U.S./French satellite, TOPEWPoseidon, launched in August 1992, uses a radar altimeter to precisely measure the shape of the earth (geodesy) as well as the changes in global sea states to help determine the ocean’s role in the earth’s climate.

Space Environment. James Van Allen used data from America’s very first satellite (Explorer I) to discover the “belts” of energetic particles encir- cling our world and which now bear his name. Subsequent Explorer space- craft allowed scientists to probe the and the earth’s magnetic field. The interaction between these phenomena was examined by the International Sun-Earth Explorer (ISEE) satellites, launched in 1977 and 1978, two in orbit around the earth and one positioned in a special orbit 16 Introduction to Space Sciences and Spacecraft Applications between the sun and the earth. Pioneer and Voyager spacecraft, sent to visit the planets, also sampled the interplanetary environment. These satellites are now on their way out of our solar system, but scientists continue to monitor their signals in hopes of gaining information on interstellar space. The sun has the greatest effect on the near-earth space environment, and a series of Orbiting Solar Observatory (OSO) spacecraft launched between 1962 and 1975 were used to study solar flares and temperature differences on the sun’s surface. The Solar Maximum Mission (SMM or SolarMax) was launched in 1980 to observe the sun during the maximum phase of its 11-year cycle. This spacecraft was the objective of a Space Shuttle rendezvous and EVA repair mission in 1984. UZysses, a joint U.S ./ spacecraft deployed from the Space Shuttle in October 1990, used a gravity assist from to send it into a solar- polar orbit to allow mapping the solar environment in this previously unexplored region of our solar system. Long-term effects of the space environment on spacecraft has been inves- tigated with the Long Duration Exposure Facility (LDEF). Released from the

Figure 1-1 1. The SolarMax sun-observing satellite was the object of a Space Shuttle repair mission in 1984. Introduction and History 17

Space Shuttle in 1984 for a one-year stay, the Challenger accident delayed its recovery until 1990 when it was finally returned to earth for analysis.

Planetary Exploration. Planetary exploration began with flights to our nearest neighbor, the moon. Before sending men there, several Ranger spacecraft were sent crashing into the lunar surface, sending back high- resolution photographs right up until impact. Lunar Orbiters pho- tographed the moon from orbit to locate possible landing sites, and Sur- veyor craft soft-landed on the moon and sent back photos and conducted lunar soil sampling experiments. Mariner 2 became the first spacecraft to fly by another planet, passing within 21,000 miles of Venus in December 1962. Subsequent Mariner and Pioneer spacecraft were sent toward Mercury, Venus, Mars, Jupiter, and Saturn-all the planets known to early astronomers! In 1976, two Viking spacecraft successfully landed on Mars, sending back pictures of the Mar- tian landscape and conducting soil and atmospheric experiments. Spec- tacular pictures and a wealth of planetary information on Jupiter, Saturn, Uranus, and Neptune were sent back from the two Voyager spacecraft, now on their way out of our solar system. The Magellan spacecraft used synthetic aperture radar (SAR) to map more than 98 percent of the surface of the cloud-enshrouded planet Venus from September 1990 to September 1992. Galileo, deployed in October 1989 from the Space Shuttle, used

~-- - _-~-

Figure 1-12. The Pioneer spacecraft made the first “up close” observations of many of the planets. 18 Introduction to Space Sciences and Spacecraft Applications both Venus and the earth for gravity assist boosts to reach the planet Jupiter and release a probe into its atmosphere. Along the way, Gulileo flew by the asteroids Gaspra (October 1991) and Ida (August 1993), giv- ing us our first close looks at these solar system companions. The Soviets sent numerous spacecraft to explore Venus, and two of their Venera craft successfully landed on the surface in March 1982. The landers were only able to transmit pictures and data for a short time until the harsh environment (temperatures which will melt lead and pressures about 100 times that on earth) disabled the spacecraft. While the United States con- ducted the first flyby of a comet (Giaccobini-Zinner) in 1985, a flotilla of spacecraft from the Soviet Union, the European Space Agency, and Japan made somewhat more noteworthy flybys of Halley’s comet in 1987.

Space Exploration. An Orbiting Astronomical Observatory (OAO 2) was launched in December 1968 with a complement of ultraviolet instruments designed to examine the stars, and discovered evidence of the existence of “black holes.” High-Energy Astronomy Observatory (HEAO) spacecraft were launched between 1977 and 1979 to conduct X-ray, gamma ray, and surveys of the universe, and the International Ultraviolet Explorer (IUE) conducted observations in the UV wavelengths. Launched in November 1989, the Cosmic Background Explorer (COBE) carried instruments designed to search the infrared background of deep space for information about the “Big Bang” theory of the origin of our universe. Deployed in April 1990 from the Space Shuttle, the Hubble Space Tele- scope (HST) was the first of the “great observatories,” facility-class satel- lites designed to offer scientists unparalleled opportunities to explore the universe. HST has a 2.4-meter (94.5-inch) diameter mirror which can pick up objects 50 times fainter and 7 times more distant than present ground- based observatories, potentially expanding the universe visible to astronomers by five hundredfold! Designed for on-orbit servicing, in December 1993 astronauts were able to insert a set of corrective optics to offset the “spherical aberration” flaw discovered in the primary mirror that compromised HST’s capabilities. The second great observatory, the Compton Gamma Ray Observatory (GRO), was deployed from the Space Shuttle in April 1992. GRO was designed to capture the relatively unim- peded gamma ray emissions from cataclysmic cosmic events like super- novas, black holes, and even the remnants of the Big Bang. Other great observatory satellites may follow. Introduction and History 19

Figure 1-13. Hubble Space Telescope. The HST was the object of a space shuttle repair mission in 1993.

Commercial Satellites. By far, the largest commercial interest in space systems is in the relay of information, or communications. Communica- tion satellites had a humble start with Echo I, launched in August 1960, which was simply a huge -reflecting balloon in low-earth orbit. By 1963, Syncom I demonstrated the capability and advantages of communi- cating via satellite in a . Today, over one hundred com- munication satellites exist in this orbit, each capable of relaying thousands of telephone calls and television pictures simultaneously around the globe. In 1962, President Kennedy signed a bill creating the Communica- tions Satellite Corporation (Comsat), a joint government-industry under- taking allowing transfer of US.-developed satellite communications tech- nology for commercial use. The International Telecommunications Satellite Consortium (), with Comsat representing the United States, promotes international use of these capabilities. Both organizations have been highly successful. A similar arrangement exists involving dissemination of data gathered by the Landsat satellites for which demand by the commercial public has grown. The French SPOT (Satellite Probatoire de I’Observation de la Terre) satellite competes in this area, offering high resolution photographs of points of interest on the earth to news media, governments, and even military organizations worldwide. 20 Introduction to Space Sciences and Spacecraft Applications

Figure 1-14. Gorizont. Commercial channels have been made available on this Russian geosynchronous .

Military Satellites. The early experiments using captured V-2 rockets were conducted under the auspices of the military and led to the develop- ment of the intercontinental ballistic missile (ICBM). As other uses of space became apparent, the military conducted their own experimental programs and developed systems to take advantage of the new “high ground” of space. Today, the military routinely conducts force enhancement missions such as communications, navigation, and remote environmental sensing using dedicated systems in space. Surveillance and monitoring missions are also routinely conducted, though the systems used are often classified. However, the worth of these systems became generally apparent during the war with Iraq, where coalition forces scored a decisive victory using these capabilities heavily.

International Use of Space

Although the foregoing pages emphasize U.S. accomplishments and activities in space, Russia, Japan, France, and China all have fully devel- oped programs capable of providing spacecraft manufacture, launch ser- vices, and operation of payloads. The scientific, environmental sensing, and commercial communications use of space systems is an international arena. The United States and Russia no longer dominate the field, with France, China, and others vying for commercial customers for space sys- Introduction and History 21

Figure 1-15. Navstar/GPS. Though designed to provide precision navigation for military forces worldwide, commercial use of the Global Positioning Satellite system has become widespread. tems and services. Nations such as the United Kingdom, Canada, Italy, India, Korea, and others have a presence in space with specific satellites or technologies geared to space applications. Space has become a world- wide area of interest and commerce with many nations joining together to pursue mutual interests. The European Space Agency (ESA), a consor- tium of space organizations from 13 countries, is one such example. A listing of many U.S. and foreign satellites is given in Appendix B.

The Future

While the Space Shuttle will continue to be used to deliver people and material to outer space, future plans call for a more robust mix of launch vehicles which may include an increasingly wider range of expendable launch vehicles (ELVs) or more reusable concepts such as follow-on space shuttle designs, single-stage-to-orbit (SSTO) vehicles, or a Nation- al Aerospace Plane (NASP) capable of reaching space using more con- ventional take-off and landing operations. The Russians have tested a shuttle-like craft, Burun (“Snowstorm”), although the likelihood of its fur- ther development and use is doubtful. The Russians also have developed a heavy-lift booster of the Saturn V class, Energiu, which could be used to launch complete space station modules into low-earth orbit. The Euro- 22 Introduction to Space Sciences and Spacecraft Applications pean Space Agency continues to upgrade the Ariane booster and provide commercial launch services to customers worldwide. The Russians continue their exploitation of space through continued use and expansion of the Mir space station, which includes long-duration missions of American astronauts and visits of the U.S. Space Shuttle. The Russians will also contribute to the international space station effort along with the United States, Canada, Japan, and the European Space Agency. Uses of the space station will include biomedical and material processing experiments conducted within the laboratory modules, as well as exter- nally-placed earth-observing and astronomical instruments. The Ballistic Missile Defense Office (BMDO-formerly the Strategic Defense Initiative, or SDI) envisions a missile defense system which includes space-borne assets. Success in the Persian Gulf War ensures con- tinuance and updating of military satellite systems including the Global Positioning System (GPS) for navigation, the MILSTAR and other net-

Figure 1-16. International space station. This 1993 concept shows the basic building blocks of the planned space station: habitation and laboratory modules supported by solar power arrays and other systems on a long truss structure. Introduction and History 23 works for communications, and other military systems for weather and surveillance uses. NASA’s Mission to Planet Earth is an effort to characterize the earth’s environment and the changes it may be undergoing. This effort includes the Earth Observing System (EOS), a series of large “observatory”-sized satellites designed for simultaneous measurements of the earth’s environ- ment from space, as well as several smaller specialized satellites and instruments from many countries. NASA’s Mission from Planet Earth looks outward to man’s continued exploration of the universe. It includes the moodMars initiative which promotes a lunar base and a manned mission to Mars, possibly as an inter- nationally cooperative effort. This initiative includes unmanned missions to these and other bodies in our solar system and provides an impetus for increased development and use of robotic systems. It is obviously man’s intention to expand his presence and capabilities in this frontier, and the future of space exploration, no matter how unpredictable, will certainly be challenging, exciting, and international in character.

REFERENCES/ADDITIONAL READING

Kerrod, R., The Illustrated History of NASA, Anniversary Edition. New York: Gallery Books, 1988. Bilstein, R., Orders of Magnitude: A History of the NACA and NASA, 1915-2990. Washington, D.C.: US.Government Printing Office, 1989. Clark, P., The Soviet Manned Space Program. New York: Orion Books, 1988. Simpson, T. (Ed.), The Space Station. New York: IEEE Press, 1985. Thompson, T. (Ed.), TRW Space Log, Vol. 27, 1957-1991. Redondo Beach: TRW Space & Technology Group, 1992.

EXERCISES

1. List and expound upon some of the benefits achieved by using remote sensors in space. 2. Repeat Exercise 1 for satellite navigation systems. 3. Repeat Exercise 1 for satellite communications systems. 24 Introduction to Space Sciences and Spacecraft Applications

4. Expound upon the beneficial uses of the above systems to units of the various military services (Army, Navy, Air Force, Marines, Coast Guard). 5. List any other applications you can think of for using systems in space. 6. Repeat Exercise 1 for one of these other applications. 7. Discuss the differences in the approaches taken in space by the Unit- ed States and the Soviet Union (RussidCommonwealth of Indepen- dent States). 8. What do you think the ramifications of other countries (or groups of countries) establishing an independent space capability might be? 9. What, in your opinion, should be the next major effort in man’s ven- ture into space (e.g., space station, moon base, Mars, commercializa- tion) and why? 10. What are your opinions about international cooperation in these efforts? PART 1 Space Sciences CHAPTER 2 Orbital Principles

Around 350 B.C., the renowned Greek philosopher Aristotle made the deduction that was to influence man’s perception and understanding of the motions of the planets for over two thousand years! Aristotle was of the opinion that all the heavenly bodies, including the sun, the wanderers (planets), and the stars, were in circular motion about a fixed and unmov- ing Earth. So authoritative was Aristotle that the Catholic church later accepted this view as its official doctrine on the matter and used its reli- gious influence to enforce this position. As a direct result, Galileo was arrested for espousing views to the contrary in 1600 A.D.! However, a century before Galileo, the first serious questioning of Aristotle’s views had already begun. Astronomy (or rather astrology) was very popular at the time of Coper- nicus, around 1500 A.D., and the inability to accurately describe the motions of the heavenly bodies increased efforts to find a better explana- tion than Aristotle’s. Copernicus was both an astronomer and mathemati- cian, and he used the observed angular positions of the planets and trigonometry to correctly place the solar system in its proper order, with the sun at the center and the earth as just another wanderer in motion about the sun like the other planets. However, the idea of a moving earth and similarity with the other planets was quite radical at the time and made his heliocentric (sun-centered) hypothesis hard to accept. For instance, Tycho Brahe, the late sixteenth-century astronomer, com- pletely rejected the notion. Tycho, in the years prior to his death in 1601, conducted the most exhaustive and accurate recording of the movements of the planets to date. He was sure that his data held the secret to the mys- tery of planetary movements, but his mathematical ability was too poor to check out his theories. Therefore, Tycho solicited the assistance of math- ematicians like Johannes Kepler. After Tycho’s death, Kepler came into possession of most of the obser- vational records kept over the many years. Kepler believed in the Coperni-

26 Orbital Principles 27 can heliocentric hypothesis but was still saddled with the religious-backed belief in the perfection of the heavens, which called for purely circular motions of the planets. For nine years, Kepler struggled to fit the observed motions of the planet Mars into different models of combinations of circu- lar paths. It was only after finally trying an oval-shaped path that Kepler found that the orbit of Mars fit extremely well into an ellipse. Kepler sum- marized his findings in his famous three “laws” which will be examined in this chapter. Kepler showed that these relationships held not only for the paths of the planets around the sun, but for the paths of the four recently discovered circling the planet Jupiter by Galileo. We will use these same relations to describe the motion of earth-orbiting satellites. Galileo Galilei was well known, even in his own time, as a capable sci- entist and experimenter in many fields. In his quest for knowledge, Galileo employed a systematic approach to his studies that laid the foun- dation for what is now known as the scientific method of investigations. His study of the motions of bodies led to the understanding of friction, inertia, and the acceleration of falling bodies. Galileo was the first to use a combination of lenses to make telescopic observations, and with these crude devices he made some astounding discoveries, including the exis- tence of the moons around Jupiter now known as the Galilean moons. It was the combination of Galileo’s experiments in general motions and Kepler’s findings in planetary motions that gave Newton the tools he needed to explain, not just describe, the mechanics of orbits. Sir Isaac Newton’s status as a genius is obvious considering the great leap of understanding in many different areas attributed to the seventeenth-centu- ry physicist. By his time, people had had over 50 years to mull over the find- ings of Kepler and others concerning planetary motions. It is important to recognize that Kepler’s “laws” were derived empirically, simply describing the characteristics of the planetary motion recorded by Brahe. Yet no one had come up with a plausible explanation as to why the planets followed the par- ticular paths around the sun that Kepler’s relationships indicated, and there was great debate among the scientific investigators of the time. Sir Edmund Halley, a friend of Newton’s, nonchalantly mused on the topic one day, and to his amazement Newton replied that he had come up with an explanation some 20 years earlier but had not bothered to publish the results. At Halley’s urging and expense, Newton published his findings which included his three laws of motion and his description of gravitational force. Newton and Kepler formulated their “laws” based on the motions of the planets around the sun, but these relationships describe orbital motions 28 Introduction to Space Sciences and Spacecraft Applications between any two bodies in the universe! In the following sections we will look at these relationships and how they describe the motions and other characteristics of bodies in orbits around the earth.

ORBITAL PRINCIPLES

Kepler’s Laws

Kepler’s First ,aw. Kepler’s first law reveals that :aptive” satellites (those with closed orbital paths) will travel around the earth in elliptical (or circular) paths with the center of the earth located at one of the foci, as depicted in Figure 2- 1.

Orbital Parameters. Figure 2-1 defines some of the terms with which we will be dealing, and the geometry of the ellipse reveals some useful relation- ships between these orbital parameters. From the figure it is obvious that:

a = semi-major axis b = semi-minor axis Note: a circle is simply an ellipse with a = b r = radial distance between bodies’ mass centers ra = apoapsis radius (maximum distance between bodies) rp = periapsis radius (minimum distance between bodies) v = true anomaly (measured in same direction as movement in the orbit)

Figure 2-1. The ellipse. Many orbital terms can be defined simply by the geometry of an ellipse. Orbital Principles 29

ra+rp=2a or a=- ra + rp 2 where ra and rp represent the upoapsis (maximum) and periapsis (mini- mum) distances between the bodies, respectively. These distances will be better defined shortly. The quantity a is defined as the semi-major axis of the ellipse. The eccentricity e describes the shape of an ellipse, sort of in terms of how fat to wide with respect to the semi-major axis and semi-minor axis b:

A more useful form for the eccentricity equation can be derived from geometry as:

ra - ‘p e=- (2 - 3) ra + rp

The orbital parameters a and e together define the size and shape of an ellipse, as exemplified in Figure 2-2 in which both ellipses have the same semi-major axis but different eccentricities. The general equation for a conic section, of which an ellipse is just one type, may be written in a form which reveals many useful relationships:

a(1 -e2) r= (2 - 4) 1 + e(cos u)

a=3 b=l a=3 b=2 e = 0.94 e = 0.75

Rgure 2-2. Eccentricity defines the shape of an ellipse. 30 Introduction to Space Sciences and Spacecraft Applications

In an orbit, r represents the radial distance (orbital radius) between the bodies’ mass centers. The true anomaly 2) is the angle measured from the major axis line (in the direction pointing toward periapsis) to the radial line between the two bodies. This is measured in the same direction as the motion of the orbiting body (refer to Figure 2-1).

Example Problem:

With the knowledge (from geometry) that for a circle the dis- tances a and b are equal, find the eccentricity of a circular orbit and give the relationship between the periapsis, apoapsis, and semi- major axis distances.

Solution:

From equation 2-2 with a = b, it is clearly seen that the eccentric- ity of any circular orbit is zero. Relating this finding to equation 2-3, for this expression to be zero, ra must be equal to rp. In fact, inserting a value of zero for eccentricity into the conic sec- tion equation (eq. 2-4) we find that r = a for any true anomaly 2). This result tells us that the orbital radius for any circular orbit remains constant throughout the orbit.

Orbital Types. A closer look at the conic section equation (eq. 2-4) reveals that the eccentricity can tell us immediately what type of orbit we are in: If we only consider positive eccentricities, the first limiting case is when e = 0, which we already know indicates a circular conic section and thus, a circular orbit. The next limiting case occurs when e = 1. For this value the conic equation gives a value of infinity for the radius at the point where the true anomaly approaches 180”. This corresponds to a parabolic conic section. Values of eccentricity greater than 1 indicate a hyperbolic conic section. Parabolic and hyperbolic orbits represent “open” or non- repeating orbits. Such orbits are used by deep space probes to escape the earth or even the solar system as in the case of the Voyager probes. Orbital Principles 31

The final case, then, is when the eccentricity is greater than zero but still less than one (0 < e e 1). For fixed values of the eccentricity and semi-major axis, we can see that the orbital radius is a function of the true anomaly which is a measure of exactly where along the conic sec- tion the orbiting body is. Substituting different values for the true anomaly, one would find that the minimum radius occurs when u = 0" and the maximum radius occurs when u = 180".These distances cor- respond to the periapsis and upoupsis points in the orbit shown in Fig- ure 2- 1 and prove that they are associated with the minimum and max- imum distances between the bodies in an elliptical orbit. Values of eccentricity within this range, then, correspond to an elliptical conic section and would indicate an elliptical orbit. At periapsis (u = 0') and at apoapsis (u = 180°), the conic section equa- tion (eq. 2-4) simplifies to reveal two useful relationships:

r, = a(l + e) and rp = a(l - e) (2-5)

It is important to realize that orbital radius increases continuously from periapsis to apoapsis, and decreases continuously from apoapsis to peri- when moving in an elliptical orbit.

Kepler's Second Law. Kepler's second law reveals that a line drawn between the two bodies will sweep out the same amount of area during the same time period anywhere along the orbital path. This characteristic is illustrated in Figure 2-3.

At2 = At1 Areal =Area2 Figure 2-3. Aredtirne relationship. Kepler's second law reveals that the velocity of an object changes with orbital radius. 32 Introduction to Space Sciences and Spacecraft Applications

The most important point this law reveals is that the speed of an object in an orbit changes with changing distance between the bodies. For the areas shown in Figure 2-3 to be the same, the orbiting body must slow down when it is farther away so that the line between the bodies sweeps out the same area in the same amount of time as when it is traveling clos- er (both At’s shown are equal). When Newton’s laws are discussed, we will find a useful expression which shows this dependency of velocity on orbital distance.

Kepler’s Third Law. In his third law, Kepler reveals that a relationship exists between the semi-major axis and the period of an orbit. Stated mathematically:

In equation 2-6, T represents the , or time it takes to travel through one complete orbit with semi-major axis a. The term p represents a gravitational parameter which has a specific value for each body around which an orbit may be described. The gravitational parameter for the earth is:

CI English units Metric units Earth 1.4077 x 10l6 ft3/sec2 3.986 x lo5 km3/sec2

Use of this form of the gravitational parameter in Kepler’s time equa- tion (eq. 2-6) with the proper units used for the semi-major axis results in an orbital period given in seconds.

Example Problem:

With the knowledge that it takes the moon 27.32 (sidereal) days to complete one orbit around the earth, determine the semi-major axis of the moon’s orbit. Note: a sidereal day is the time it takes for the earth to complete one full rotation about its axis with respect to the (inertially fixed) stars. A solar day, measured with respect to the sun, differs due to the fact that the earth is also in motion around the sun.

1 sidereal day = 23 hrs, 56 min, 4 sec = 86,164 sec Orbital Principles 33

Solution:

T = (27.32 days) x (86,164 sedday) = 2,354,000 sec

Solving equation 2-6 for the semi-major axis:

a3 = (4)~2 (2 - 7) 4n:

we find that for the period computed above:

a = 1.26 x lo9 ft = 237,658.8 mi (using 5,280 ft/mi) = 382,469.6 km (using the metric values)

The semi-major axis represents the mean distance of an orbiting body from the center of the earth. If the moon’s orbit were circular, this distance from the earth would remain constant (as we have already shown). In fact, the moon’s orbit is slightly elliptical, so its distance from the earth is con- stantly changing as the following problem shows. Note: the apoapsis and periapsis points of an orbit described around the earth are called apogee and perigee respectively.

Example Problem:

The moon is in a slightly elliptical orbit around the earth (e = 0.055). Determine the moon’s apogee and perigee distances in both miles and kilometers.

Solution:

Recalling the values obtained earlier for the moon’s semi-major axis, and using the relationships obtained from the conic equation at apoapsis and periapsis (eq. 2-5):

r, = a(l + e) = 250,730.0 mi = 403,505.4 km

rp = a( 1 - e) = 224,587.6 mi = 361,433.7 km 34 Introduction to Space Sciences and Spacecraft Applications

This represents an appreciable change in distance for an orbit considered almost circular and would certainly have to be taken into account if plan- ning a trip to the moon.

Newton’s Laws

Kepler’s laws were empirically derived relationships describing plane- tary motions based on years of observations. Newton’s laws more pre- cisely defined the mechanics of motions and, combined with his idea of universal gravitation, allowed a more complete description of general orbital motions. In a dedicated course in orbital mechanics, Newton’s laws of motions are studied in great depth, and in fact, Kepler’s laws are usu- ally derived from these relationships. For our purposes, the next few sec- tions will merely state some of the important results of Newton’s laws.

Angular Momentum. The angular momentum of an orbit is represented by a vector quantity perpendicular to the orbital plane (the plane contain- ing the satellite position and velocity vectors) as shown in Figure 2-4. One of the results of Newton’s laws shows that the angular momentum of a satellite’s orbit is constant, which for a vector means constant in both magnitude and direction. This implies that an orbit lies in a plane which remains inertially fixed in space. Consider, in Figure 2-5, that the vector h shown represents the angular momentum of a satellite in orbit around the earth. (We are looking edge- on to the satellite’s orbit so its orbital plane looks like just a line across the

satellite - orbit position h angular vector

satellite velocity vector

Figure 2-4. The angular momentum vector of an orbit is perpendicular to the orbital plane. Orbital Principles 35

- S h

orbit

Figure 2-5. Constant angular momentum. This property of an orbit results in an inertially-fixed orbital plane. earth in the figure.) The figure shows that in half a year, as the earth has moved from one side of the sun to the other, the angular momentum (and thus the plane) of the satellite’s orbit has remained pointing toward the same inertial direction in space. During this time, the earth has also been rotating daily beneath the satellite’s fixed orbital plane. This affects the satellite’s ground track (the path which the satellite appears to etch across the earth’s surface), a point which we will discuss shortly.

Total Energy. Another important result of Newton’s laws states that the total energy of an orbiting body is also a constant. The expression for the specific total energy E of an orbiting body is:

E=-v2 --CL 2r

The first term on the right-hand side of the equation can be seen to relate to the kinetic energy of the orbiting body and the second term rep- resents the body’s potential energy. For the total energy of an orbiting body to remain constant, the kinetic energy term must decrease as the potential energy term increases. (Since the potential energy term is a neg- ative value in the equation above, this term will increase [become less negative] as the orbital radius r increases.) This corresponds to what was stated earlier in Kepler’s second law and demonstrated in Figure 2-3. 36 Introduction to Space Sciences and Spacecraft Applications

Orbital Velocities. Another very useful expression for the total energy is:

Relating this to equation 2-8, we can write:

--I--V2 - -cL 2 r 2a and solving for velocity we get:

(2 - 10)

Equation 2- 10 is a general equation for velocity at any point in an orbit based on its orbital radius r. This equation can be simplified for three spe- cific cases of particular interest: In a circular orbit we know that r = a always, so equation 2-10 may be rewritten as:

(2-11)

This simple relationship shows that circular orbital velocities vCimare constant and determined by the circular orbital radius rcimonly. From equation 2-5, we know that at apoapsis in an elliptical orbit, r = ra = a(l + e). Using this in equation 2-10, we can write:

(2 - 12) va=J= a(l +e)

Since this corresponds to the point of maximum orbital distance, Kepler and Newton’s laws tell us that this apogee velocity v, is the slowest during the orbit.

Similarly, using r = rp = a(1 - e) for the orbital radius at periapsis, equation 2-10 becomes: Orbital Principles 37

vp =J- (2- 13)

Being associated with the minimum orbital radius, the periapsis velocity vp is thefastest velocity in the orbit.

We now have enough information to analyze established orbits if given just a few of the orbital parameters.

Example Problem:

Due to thrust limitations and the reaches of the atmosphere, the Space Shuttle is limited to operations between about 200 km to 800 km altitudes. From this information, determine the orbital parame- ters associated with an elliptical orbit between these two altitudes.

Solution:

a. Semi-major axis (equation 2-1): a = 6,878 km

Note: altitudes must be converted to orbital radii by adding the radius of the earth, & = 6,378 km for use in equation 2-1.

b. Orbital period (equation 2-6): T = 5,677 sec = 94.6 min

Slightly over an hour and a half between sunrises!

c. Total specific energy (equation 2-9): E = -28.98 km2/sec2

Note: Like eccentricity, knowledge of the total energy can indicate the type of orbit as well. All “closed” (elliptical and circular) orbits have negative values for total energy. Zero or positive total energies indi- cate the “open” (parabolic or hyperbolic) orbits introduced earlier.

d. Eccentricity (using equation 2-3 or equation 2-5): e = 0.0436 38 Introduction to Space Sciences and Spacecraft Applications

This value (between 0 and 1) indicates an elliptical orbit and agrees with the closed orbit indicated by the negative value obtained for the total energy.

e. Apogee velocity (equation 2-12): v, = 7.29 Wsec = 26,235 km/hr

f. Perigee velocity (equation 2-13): vp = 7.95 Msec = 28,628 km/hr

These are appreciable velocities and the velocity difference between apogee and perigee is also not insignificant!

ORBITAL ELEMENTS

To completely describe the shape and orientation of an orbit around the earth, six quantities must be specified as shown in Figure 2-6. During the following discussion keep in mind that, once set, the plane in which an orbit lies remains inertially fixed in space. As stated earlier, the eccentricity e and semi-major axis a define the shape and size, respectively, of the orbit. The orientation ofthe orbital plane is described by two angles: the inclination i and the longitude of the ascending node $2.The inclination describes the angle that the orbital plane makes with a reference plane (chosen to be the plane of the equator for earth-orbiting satellites). This is the same as the angle between the angular momentum vector of the orbit h and a coordinate axis perpendicular to the reference plane (K chosen in the direction of the North Pole) as shown in Figure 2-6. The longitude of the ascending node is an angle that describes the rota- tion of the orbital plane from a line within the reference plane which points toward an inertially fixed direction (chosen toward the direction of the vernal equinox, an essentially inertially fixed direction infinitely far away in space) to the line formed by the intersection of the reference and orbital planes (line of nodes) on the side where the orbital motion is from south to north (“ascending” node). The argument of perigee o describes the orientation of the elliptically- shaped orbit within the orbital plane. It is measured as the angle from the line of nodes to the radius of perigee of the orbit. Orbital Principles 39

1 st pt. (vernal

Figure 2-6. Orbital elements. Six independent quantities are required to completely describe an orbit.

Finally, a parameter must be used to describe the actual position of the orbiting body at any particular time or location within the orbit. Either an angular measurement, such as the true anomaly v, or a time parameter, such as time since perigee passage T,, may be specified.

ORBITAL PROPERTIES

This section discusses some of the properties of a satellite due to its unique orbital position above the earth.

Field of View

Perhaps the most obvious benefit of being in space is the view one has when looking back at the earth. The amount of the earth in view at any one time is a function of the height above the surface as illustrated in Fig- ure 2-7. In this figure, Rerepresents the mean radius of the earth (6,378 km or 3,444 nm), and h is the height of a satellite above this mean radius. The point directly beneath the satellite on the surface of the earth is known as the nadi,: The angle Cp is known as the earth angle and represents the maximum angular portion of the earth, measured to the tangential horizon, visible from the satellite height. The equation for determining Cp for any given altitude is: 40 Introduction to Space Sciences and Spacecraft Applications

h=35,786 krn

Figure 2-7. Field of view. Properties that a spacecraft has with respect to the surface of the earth change dramatically with altitude.

q=cos-l ~ (2 - 14) (R:: (R:: h)

The maximum distance visible from one tangential horizon to the other is known as the swath width SW which can be found from:

SW = 2R& (2-15)

The area on the surface of the earth associated with the swath width is shown shaded in Figure 2-7. This is thefootprint area of coverage theo- retically possible from the satellite position. This area can be found from:

A = 2nR; sin@ (2-16)

Finally, the angle subtended (from the perspective of the satellite) between the line from nadir to a line toward the tangential horizon is known as the angular field of view (FOV) Q and is determined from:

(2 - 17)

This angle is important, as it represents the angle that a particular sensor would have to be able to scan or view in order to cover all the potentially visible surface of the earth. All of these terms have similar but slightly different meanings when used to describe the capabilities of a particular sensor in space, as dis- cussed in Chapter 6. For instance, swath width may be used to describe Orbital Principles 41

only the side-to-side distance over the surface of the earth scanned by the sensor, which may be much less than the geometric swath width described here. The same is true of the footprint area viewed by a particular sensor. When discussing the angles through which a sensor may scan, the sensor field of view may be different from the angular field of view described here, and the angles viewed may also be identified as the nadir angle or the look angle. Usually, the context in which these terms are used will clarify the intended meaning.

Ground Track

As mentioned earlier, the ground truck represents the path of the satel- lite superimposed onto the surface of the earth. This path is a result of a combination of the orbital motion of the satellite and the movement of the earth due to its rotation. Figure 2-8 shows the ground track of a typical low-earth orbiting satellite through four orbits. The orbit is circular with an altitude of about 1,000 km in an orbital plane inclined about 45" from the equator. The most notable feature of such a ground track is the apparent move- ment of the orbital path toward the west. As we learned earlier, the plane in which an orbit lies is inertially fixed in space. The apparent westward movement of the ground track is caused by the eastward rotation of the

Figure 2-8. Ground track. The path an orbiting satellite traces across the surface of the earth changes due to the earth's rotation. 42 Introduction to Space Sciences and Spacecraft Applications earth beneath the orbital plane. The distance between the first ascending (south to north) passage through the equator shown in Figure 2-8 (1) and the next ascending passage (2) is determined by the period of the orbit which, as we saw in equation 2-6, is a function of the orbital semi-major axis only. Another point to notice in Figure 2-8 is that the ground track is con- strained between about 45" north and 45" south latitudes. This corresponds directly to the , and constrains the ground track only; the area on the surface of the earth visible to the satellite is a function of the height of the satellite within the orbital plane as mentioned earlier.

Maximum Time in View

The footprint of the satellite's view on the earth moves continuously with the satellite as it travels around the earth. This means that the satel- lite can only view a particular spot on the surface (or conversely, a ground observer could only see an orbiting satellite) for a limited amount of time. This amount of time depends on the satellite's altitude (which corresponds to the satellite velocity) and orbital inclination, as well as on the latitude and distance from nadir of the ground observer during the satellite pass. For planning purposes, the maximum time that a satellite can be in view, associated with tangential horizon to tangential horizon passage directly over the observer's location, can be found from:

(2-18)

Note that use of this equation requires computation of the arc cosine in degrees. Of course, any oblique pass not directly over the observer's loca- tion will result in less time that the satellite will be in view.

Number of Revolutions per Day

The number of times a day that a satellite will completely circle the globe depends on its orbital period, which is a function of the semi-major Orbital Principles 43 axis, which can be related to the satellite altitude. For circular orbits (only), a simple expression gives the number of revolutions per day as:

## Revlday = 16.997 ~ (2 - 19) (Rri h)li2

Revisit Time

It may be desirable to have a satellite return over a specific point on the earth's surface periodically for monitoring purposes. It may even be desir- able to have the satellite pass over the specified point at the same time of day each visit, or remain in view of a spot indefinitely. Using combina- tions of the orbital elements and parameters, it is possible to create an orbit to meet many of these requirements; however, there is no simple way of describing this process here. Nonetheless, for some simple cases such as discussed in the following sections, it is possible to see how some of these requirements can be approached. Revisit times are mainly dependent on orbital period T which, as we saw, is a function of the semi-major axis of the orbit. If an orbital period is established which divides evenly into the period of one sidereal earth rota- tion (T, = 1 sidereal day = 86,164 sec) such that TJT = n where n is an integer, then after n orbits (neglecting any perturbations) the ground track of the orbit will begin to repeat. If n turns out to be unity, then the orbital period exactly matches the rotational period of the earth and the ground track repeats itself each orbit. This is known as a .

Example Problem:

The lowest altitude at which a satellite may circle the earth in a repeating orbit (though its lifetime would be limited due to atmos- pheric drag) is 262 km. Determine the footprint angle, swath width, footprint area, angular field of view, maximum time in view, and the number of revolutions per day. Show that the satellite period is an multiple of the sidereal day.

Solution (using equations 2-14 through 2-19 and equation 2-6):

Cp = 16.15" = 0.28 rad SW = 3,595.3 km (use radians for Cp in this equation!) 44 Introduction to Space Sciences and Spacecraft Applications

A = 71,089,195.3 km2 !2 = 73.85' = 1.29 rad tview= 8.59 min (not a lot of time if you have much to com- municate between the satellite and a ground station!) # Revlday = 16 T = 5,384.7 sec = 89.75 min T,/T = 16

USEFUL ORBITS

There are many orbits which have come into common use. This section presents four of the most common.

Low-Earth Orbii

There is no specified cut-off altitude, but as the name implies, low- earth orbits (sometimes designated LEO) represent orbits which come rel- atively close to the surface of the planet. All the manned spacecraft, with the exception of the Apollo missions to the moon, have been in low-earth orbits, as are many of the earth-observing satellite systems. As the previ- ous sections have implied, low-earth orbits are characterized by short orbital periods, many revolutions per day, high orbital velocities, and lim- ited swath areas on the earth's surface.

Geostationary Orbit

In the last section, an orbit was defined which had an orbital period exactly equal to one sidereal day. Any orbit with this period can be called geosynchronous. A geostationary orbit is defined as a circulal; geosyn- chronous orbit with an inclination equal to zero (equatorial). A satellite placed in such an orbit would appear, to an observer on the surface of the earth, to remain stationary in the sky. The higher altitude in Figure 2-7 shows the approximate position of a geostationary satellite and its associ- ated field of view. Geometrically, a satellite at geostationary altitude can see a footprint about 70 degrees north and south of the equator. The benefits of a satellite so positioned are obvious. A single sensor placed at a longitude equal to the center of the United States could moni- tor the weather pattern for the entire country continuously. A communi- cations relay station placed at a longitude in the middle of the Atlantic or Orbital Principles 45

Pacific oceans could allow direct communications between countries on either side or ships upon the seas. Many more uses exist, and the popu- larity of this operating position is producing a crowded area in space.

Polar Orbits

Strictly defined, polar orbits are orbits with an inclination of 90” which would pass over the earth’s poles each orbit. In actual use, many orbits with inclinations near 90” which pass over earth’s higher latitudes are also called polar orbits. The benefit of such an orbit, besides being able to view the higher latitudes, comes from the fact that the orbital plane is fixed and the earth rotates continuously beneath this plane. If the orbital period is a nonintegral multiple of the sidereal day, then eventually all areas of the globe will pass beneath the orbiting satellite. Of course, integral multiples of the sidereal day can allow a satellite to pass over the same point on the surface at regular intervals. This type of orbit is used regularly by satellites gathering information about the earth and its environment and resources.

Sun-synchronous Orbits

The sun-synchronous orbit can be considered a special type of polar orbit. Our discussion of orbits so far has ignored the small, but present, perturbing effects that affect the motion of an orbiting satellite. In most cases, these perturbations are accepted or simply compensated for by small station-keeping propulsion systems on board the spacecraft. In the case of a sun-synchronous satellite, though, one of these perturbing forces is used to produce a desired effect. As depicted in Figure 2-9 (greatly exaggerated), the earth is not a per- fect sphere, having more mass at the equator than at the poles. For a satel- lite in an orbit inclined greater than 0” but less than 90”, this situation imparts a force that attempts to “pull” the orbital plane toward the equa- tor. Due to gyroscopic properties, though, this force results not in a change in orbital inclination, but in a precession of the orbital plane around the equator as indicated in the figure. The effect of the perturbing force depends on the satellite altitude and the inclination of the orbit. A combination of these two can be arranged to produce a rotation of the orbital plane around the equator equal to about 46 Introduction to Space Sciences and Spacecraft Applications

force

.....-* ...... '.*. .,*-a . ..- ......

Figure 2-9. Sun-synchronous orbit. The oblateness of the earth causes a rotation of the orbital plane. one degree per day. This would mean that an orbit established perpendic- ular to the direction of the sun could be made to maintain this orientation relative to the sun throughout the year. (Refer back to Figure 2-5 which explained the fixed nature of orbital planes.) If a satellite were placed in such an orbit at an altitude producing a period that was an integral multi- ple of the sidereal day, then that satellite could be made to pass over the same spot on the earth at the same solar time(s) each day, every day of the year. Many remote sensing satellites are placed into such an orbit so as to be able to control the conditions under which the information about an area of interest is gathered throughout the year.

Constellations

As stated earlier, using combinations of the orbital elements and para- meters, an orbit can be designed to give a desired coverage of the earth. But there are purposes which require that the entire earth be covered at all times, which is clearly impossible for a single satellite. In these cases, a constellation of satellites is necessary. A constellation is simply a series of two or more satellites purposely placed into mutually supporting orbits. The satellites may be placed within the same orbital plane to effectively decrease the functional orbital period, or satellites may be placed into dif- ferent orbital planes to compensate for the movement of the observation area due to the rotation of the earth. Figure 2- 10 illustrates a constellation employing both these options (three planes of three satellites each). Orbital Principles 47

Figure 2-1 0. Satellite constellation. Multiple satellites can be placed into mutually supporting orbits.

ORBIT ESTABLISHMENT AND ORBITAL MANEUVERS

This section is included to give an orbital mechanics view of orbit establishment and orbital maneuvers. This topic is discussed again in Chapter 3 which deals with the propulsion concepts needed to perform these maneuvers.

Orbit Establishment

It was stated earlier that the total specific energy of an object in orbit was constant and could be determined from equation 2-8, which is repeat- ed here for convenience.

For spacecraft, this energy is usually established by having a launch system carry the payload to a specific altitude r = & + h with a specific velocity v before burning out. It must be remembered that the velocity is a vector quantity having both magnitude and direction. If the launch vehi- cle establishes a velocity magnitude and direction that corresponds exact- 48 Introduction to Space Sciences and Spacecraft Applications ly with the conditions for the desired orbit at that point, then the desired orbit is achieved. If these conditions are not met, some kind of orbital maneuver may be required.

Orbital Maneuvers

The basic purpose of the launch system is to carry a payload from the surface of the earth into space. However, in most cases, the launch system may not place the payload exactly into the desired orbit or may require more than just the launcher stages to achieve the desired orbit. Also, after establishing an orbit, adjustments may be required so that the orbit will conform with some desired orbital characteristic, such as period or eccen- tricity, to modify the present orbit to a completely new orbit, or merely to compensate for perturbing forces to maintain the desired orbit. Once again, as with orbit establishment, it is simply a matter of ensur- ing that the spacecraft has the proper velocity (both magnitude and direc- tion) at the proper orbital radius at the proper time. This is usually accom- plished by adjusting the spacecraft’s orbital velocity magnitude or direction until the prescribed conditions are met. This is where the term v (Av) comes from when discussing orbital maneuvers. Orbit adjustments generally fall into two types: in-plane velocity adjustments and plane changes. Addressing all possible combinations of maneuvers between different types of orbits is beyond the intent of this book; however, some basic maneuvers shall be presented to illustrate the general process.

Single-Impulse Maneuvers. Most orbital maneuvers can be considered impulsive (instantaneous) because the actual propulsion system “burn” is relatively short when compared to the typical orbital period. This allows us to assume that the change in velocity during an orbital maneuver occurs at a single point. This also tells us that the initial orbit and final orbit share at least the point at which the maneuver is performed. If the initial and desired orbits share at least one point, then a single-impulse maneuver is all that is required to change from one orbit to the other. The Av required, which may be an in-plane adjustment, a plane change, or both, can easily be computed using the velocities (magnitude and direction) of each orbit at the common point. If the initial orbit and the desired orbit do not intersect, it will not be possible to transfer between orbits with a single maneuver. Instead, an Orbital Principles 49 intermediate (transfer) orbit that intersects the initial and desired orbits will have to be chosen and multiple Av’s performed at these intersections. However, each of these Av’s may still be treated as separate single- impulse maneuvers.

Hohmann Tkansfer. A commonly used orbital maneuver between nonin- tersecting orbits is the Hohmann transfer. The classic Hohmann transfer is characterized by an elliptical transfer orbit between two co-planar circu- lar orbits as shown in Figure 2- 11. The elliptical transfer orbit also lies in the same plane as the other orbits with its perigee at the same altitude as the low orbit and its apogee at the higher orbit altitude. At the intersections (perigee and apogee) of the transfer orbit and the circular orbits, the orbital velocities are parallel and in the same direction. The Av’s are simply the differences in the velocity magnitudes at these points, with no change in direction of the velocity vectors required. The change in velocity required to enter the transfer orbit from the low circular orbit can be found from the relationships already discussed:

- orbit

Figure 2-1 1. Hohmann Transfer. This is a common method for transferring between two co-planar circular orbits. 50 Introduction to Space Sciences and Spacecraft Applications

- - AvlHobinann - 'Ptransfer orbit vcirclow orbit (2-20) At the transfer orbit apogee, the Av required to enter the circular higher orbit would be:

(2-21)

Simple Plane Changes. A simple plane change only changes the inclina- tion of the orbit and does not affect the other orbital characteristics. Plane changes must also be done at the intersection of the initial orbit and the desired orbit, as depicted in Figure 2-12. The Av required for a simple plane change can be found from:

Ai Avsirnple plane change = 2v sin (2 - 22) -2 where v represents the orbital velocity and Ai is the change in inclination. An important point that should be realized from equation 2-22 is that plane changes performed at low orbital velocities minimize the required Av. This suggests that plane changes should be performed at high altitudes or at points of lowest velocity. It should also be realized that plane changes in general are expensive in terms of Av. Equation 2-22 shows that a plane change of 60" requires a Av equal to the orbital velocity itself!

initial orbit

*"applied at intersection of orbits f%l = IT*!

Figure 2-i2. Plane change. Plane change maneuvers are performed over the equator where the two orbital planes intersect. Orbital Principles 51

REFERENCES/ADDITIONAL READING

Boorstin, D., The Discoverers. New York: Vintage Books, 1985. Bate, R., Mueller, D., and White, J., Fundamentals of Astrodynarnics. New York: Dover Publications, Inc., 1971. Eisele, J., and Nichols, S., Orbital Mechanics of General-Coverage Satel- lites. Naval Research Laboratory Report 7975, Washington, D.C.: U.S. Government Printing Office, 1976.

EXERCISES

1. Due to drag considerations, the lowest altitude at which the Space Shut- tle can operate is around 200 km. For a shuttle in a circular orbit at this altitude, determine the following orbital parameters: a. semi-major axis b. orbital period c. total energy d. eccentricity e. orbital velocity 2. Due to thrust considerations, the highest altitude at which the Space Shuttle can operate is around 800 km. For a shuttle in a circular orbit at this altitude, determine the same orbital parameters as above. 3. Compare each answer obtained in the above problems. Discuss the dif- ferences in the answers and the reasons for the differences. 4. On February 20, 1962, John Glenn became the first American to orbit the earth. At apogee, his altitude reached 260.9 km and his perigee alti- tude was 160.9 km. Determine the following parameters for his orbit: a. semi-major axis b. orbital period c. total energy d. eccentricity e. perigee velocity f. apogee velocity g. the maximum time that Glenn would be in communication range (in view) of a ground station at any time during the orbit (this would be at apogee) 52 Introduction to Space Sciences and Spacecraft Applications

5. Determine the following properties for a circular geosynchronous orbit: a. orbital period b. semi-major axis c. total energy d. eccentricity e. orbital velocity f. # revlday 6. Determine the following properties for a geostationary orbit: a) altitude b) footprint angle c) swath width d) footprint area e) field of view f) inclination g) what about the other orbital elements $2, o,@, etc.)? CHAPTER 3 Propulsion

When Newton theorized the possibility of placing an object into an orbit around the earth, not only could very few imagine a purpose for such an endeavor, no one could imagine a way to even attempt the feat. While rockets had been a part of man’s history ever since the ancient Chinese had used them for dazzling fireworks displays, uses for such systems had been limited to the purpose (primarily military) of hurling munitions from one point to impact another on the surface of the earth. In fact, this was still the main purpose even after technology allowed man to consider Newton’s musings possible. The modern ideas of rocket propulsion actually had their beginnings in the 1880s in the small Russian town of Kaluga, south of Moscow. There, Konstantin Tsiolkovsky worked out the fundamental laws of rocket propulsion and published his work proving the feasibility of achieving orbital velocities via rockets in 1903-the same year of the Wright broth- ers’ first successful powered flight! Tsiolkovsky had earlier accurately described the phenomenon of weightlessness in space (1 883), predicted earth satellites (1895), and suggested the use of liquid hydrogen and oxy- gen as propellants, which three-quarters of a century later were used to send men to the moon. With no knowledge of the work of the Russian scientist, Robert H. Goddard began studying rocketry shortly before World War I. In 1919, the Massachusetts physics teacher sent a 69-page treatise to the Smithsonian Institution entitled “A Method of Reaching Extreme Altitudes.” The “extreme altitude” he was referring to turned out to be none other than the moon, and his treatise earned him a great deal of ridicule from the press at the time. On March 16, 1926, Goddard launched the first liquid-fueled rocket in history which burned for 2.5 seconds and landed a few hundred feet away in his aunt’s cabbage patch. The press, once again, had a good time with headlines like “Moon Rocket Misses Target by 238,7991.4 Miles.” Nonetheless, Goddard, backed by grants from the Guggenheim

53 54 Introduction to Space Sciences and Spacecraft Applications

Foundation arranged by pioneer aviator Charles Lindbergh, continued his work with liquid-fueled rockets (in seclusion from the press), solving many of the practical problems involved with such an endeavor. The third rocket pioneer of note was Hermann Oberth, a German schoolteacher, whose paper “The Rocket into Interplanetary Space” (1923) gained him the attention of his country and whose research culmi- nated in the development of the V-2 rocket used by Hitler to terrorize Eng- land during the second world war. Also instrumental in the development of the V-2 was Wernher von Braun, who surrendered to American troops in 1945 and went on to become a driving force in U.S. rocket development efforts. Von Braun’s first U.S. rocket, the Redstone (which later carried America’s first man into space), was ready for launch in 1956 but, for military and political reasons, was not allowed to launch until much later. Nonetheless, von Braun was instrumental in the development of the Jupiter C rocket that placed America’s first satellite, , into orbit in January 1958 and the development of the Saturn family of rockets which were the mainstay of the Apollo program’s goal to reach the moon. But it was early in the morning of October 4,1957, that the “Space Age” truly began when the first beeps from the Soviet spacecraft Sputnik I could be heard around the globe. In those Cold War days, America was alarmed by the obvious potential this represented for a power possessing nuclear capabilities. The main U.S. space effort at that time was the Navy’s Pro- ject , which was suddenly pressured to match the Soviet achievement. The attempt proved premature as the rocket blew up on the pad on December 6, 1957, in front of a television audience of millions. Although Vanguard was eventually successful, it was the Army’s Explor- er program, headed by von Braun, which brought America into the Space Age and began the efforts, fueled by national competition, to create larg- er and more capable rockets for purposes of launching machine and man into the environment of space.

FIRST PRINCIPLES

Whatever space mission is undertaken, the spacecraft must first be put into an orbit and, secondly, may need to maneuver as well. For this it requires some sort of propulsion system and, in most cases, more than one. While there are a number of advanced types in development, the vast majority of propulsion systems today produce thrust by the expulsion of Propulsion 55 chemical propellants through nozzles. Processes for imparting energy to the propellant include chemical combustion, nuclear, electric, and others, and the lists of available propellants is also long and varied. Before we can look at these points, the principles of producing thrust via the expulsion of mass must first be understood.

The Rocket Equation

Newton’s second law of motion was discussed in the previous chapter. The law states that an applied force will result in a change in a body% momentum. Stated mathematically:

where momentum is defined as the product of mass and velocity: p = mV. The dot over the momentum term in equation 3-1 indicates the time rate of change of that quantity, and the line over a term indicates that it is a vector quantity having both magnitude and direction. In a situation where the mass does not change, equation 3-1 reduces to the familiar form F = mii, where ai represents the acceleration of the mass. However, as a rocket ejects exhaust gases for propulsion, the mass term is not constant. To determine the change in momentum, and thus the associ- ated force, we must look at the situation at two different instances in time as shown in Figure 3- 1. The figure illustrates a rocket at time t having a velocity V, and the same rocket at time t + At with a new velocity B + AV having exhausted an amount of propellant Am, which now has its own velocity Vo. If we con- sider a closed system consisting of both the rocket and the ejected pro-

m Am m- Am

d +- ...... - V VO V+ A? b t t+ At time

Figure 3-1. Rocket momentum. A rocket trades momentum with its exhausted propellant to produce thrust and increase velocity. 56 Zntroduction to Space Sciences and Spacecraft Applications pellants at all times, momentum must be conserved from one time instant to the next. At time t the momentum is:

and at time At:

The change in momentum is: - Ap = ir,t + At) - P(t) = mV + mAT - AmV - AmAv + AmVo - mV which, if we ignore the product of two differential terms, reduces to:

Ap = mAV + (50 - V)Am

Dividing through by At and taking the limit as At approaches zero:

dp dT - dm -= m- + (8, - v) - dt dt dt or, rewriting:

d8 dj5 dm m- = -+(8-T0)- (3 - 2) dt dt dt

Looking at each term of equation 3-2, we see that the term on the left-hand side of the equation represents the instantaneous ucceZerution of the rock- et mass m. The first term on the right-hand side of the equation represents the change of momentum for the entire system. Since we are considering a closed system consisting of the rocket and exhaust gases at all times, a change in the overall system momentum can only be caused by forces external to the described system. This term represents external forces (Fat) such as gravity, drag, solar pressure, and others. The vector nomen- clature shows that force direction must be understood. Some of these Propulsion 57 forces will be opposite the direction of motion, such as atmospheric drag, but some may have positive components in the direction of intended motion as can happen with solar pressure. Another external force, due to a difference between the propellant exhaust pressure and the external ambient pressure, may also exist as we will soon see. The remaining term represents the change in momentum of the exhaust gases. This exchange of momentum is the main contributor to the accelera- tion of the rocket and this last quantity is known as the thrust term (T).Note that the (vector) quantity Vo - 5 would be negative since the exhaust gasses are ejected opposite the direction of the original velocity. Rewritten as B - V,,, this quantity describes the velocity of the exhaust gasses relative to the rock- et itself. We can define this term as the exhaust velocity %. With the definition of terms given above, equation 3- 1 can be rewritten as: dij - m- = F,,, + ?; (3 - 3) dt which, in this simple form, is known as the rocket equation.

Rocket Thrust

Looking more closely at the results of the last section, it can be seen that the thrust term of the rocket equation is proportional to both the pro- pellant exhaust velocity (f - To = V,,) and the mass flow rate of propel- lant (dddt = m). We can rewrite the thrust term to better show this dependency:

To increase the thrust of a rocket then, one could try to increase either the exhaust velocity of the propellant or the mass flow rate of propellant through the rocket. To see how these quantities can be changed we must consider the characteristics of a typical rocket system such as that shown in Figure 3-2.

Mass Flow Rate. A rocket differs from a jet engine in that a rocket must carry its own oxidizer as well as fuel supply, although there are some sys- tems which simply use a single (mono) propellant. In many propulsion systems, liquid propellants are used and are delivered to the combustion chamber by mechanical pumps as depicted in the figure. The pumps con- 58 Introduction to Space Sciences and Spacecraft Applications

n

- NOZZLE P 00

Figure 3-2.Sketch of a typical rocket. Propellants are mixed in the combustion chamber and accelerated through the nozzle to produce thrust. tribute directly to the rocket thrust by controlling the mass flow rate of the propellant. Lower-thrust systems may use a pressurized bladder or gravi- ty-feed to deliver propellants to the combustion chamber. In solid rocket motors, the solid fuel and oxidizer materials are premixed and loaded into the motor casing which also serves as the combustion chamber when the fuels are ignited. Mass flow rate in solids is established by controlling the burning area and, thus, the combustion rate of the fuel. There are some obvious limitations to increasing thrust by increasing the mass flow rate. For instance, you could bum all the fuel at once, assuming you could design a combustion chamber and nozzle to handle the amount of propellant and exhaust involved. However, if you didn't just blow up, the instantaneous thrust would probably produce an unac- ceptable acceleration in the view of the structural designer or payload astronaut. More practically, the size and structural capabilities of the pumps in liquid-fueled systems limit the mass flow rates achievable. Pumps also contribute to thrust by affecting the exhaust velocity, as dis- cussed next.

Exhaust Velocity. The propellants are combined, and if necessary ignit- ed, within the combustion chamber where they create a high temperature (TJ,high pressure (p,) mixture. The velocity of the propellants in the combustion chamber is relatively slow and can be considered essentially zero, but the mixture expands to supersonic speeds through the conver- Propulsion 59

gent-divergent nozzle, leaving with an exit velocity v, at an exhaust pres- sure p,. The nozzle exit area is represented by A, in Figure 3-2. Aerodynamic and thermodynamic relationships relate the conditions between the combustion chamber and the exit area of the expansion noz- zle. Though the derivation is beyond the scope of this introduction, the exhaust velocity can be found to be:

An examination of this relationship tells us what is necessary to have a high propellant exhaust velocity, and thus, higher thrust. The term y is the ratio of the specific heats of the propellant mixture:

y=' C C"

where: cp= specific heat (under constant pressure) c, = specific heat (under constant volume)

The particular value for y depends on the properties of the particular propellant, but its value is usually between 1.2 to 1.4 and for our purpos- es can be simply considered a constant in equation 3-5. The term R represents the specific gas constant for the particular pro- pellant used. This term is actually the Universal Gas Constant (R = 8.314 joule/mole OK) divided by the molecular weight of the propellant (M kg/mole). From equation 3-5 we see that a propellant with a lower mole- cular weight will contribute to a higher exhaust velocity. This shows the benefit of using hydrogen as a propellant, as first realized by Tsiolkovsky, over heavier fuels like hydrocarbons. A higher combustion temperature To increases the exhaust velocity directly. This temperature is known as the adiabatic flame temperature and is primarily a function of the propellants used and their combustion properties. Practical limitations in combustion temperatures exist due to the structural properties of the combustion chamber and rocket nozzles. Liquid oxygen (LOX) and liquid hydrogen are kept at very low tempera- tures, and many liquid-fueled systems circulate these propellants through 60 Introduction to Space Sciences and Spacecraft Applications tubes surrounding the combustion chambers and nozzles to keep their temperatures within acceptable limits. Inside the brackets of equation 3-5 is a ratio of exhaust pressure to com- bustion chamber pressure. Minimizing this ratio contributes to higher exhaust velocities which calls for as large a combustion chamber pressure as possible. This pressure is produced by the combustion of the propel- lants; however, as was mentioned earlier, the pumps in a liquid-fueled sys- tem affect the combustion chamber pressure as well as the mass flow rate in contributing to rocket thrust. Again, limitations to achievable combus- tion chamber pressures are structural in nature.

Specific Impulse. We can define the term specific impulse (IFp!as the thrust produced per time rate of change of weight. Remembermg that thrust can be related to exhaust velocity, we can write:

The specific impulse changes for different propellants and different rocket designs and their computation is quite complicated. However, spe- cific impulse is the most commonly used criteria for comparing rocket systems since the resulting units are merely seconds. Table 3-1 shows some of the average properties of three propellant combinations.

Table 3-1 FueVOxidizer Properties

FueVoxidizer To (OK) M (glmole) I,, (sea Keroseneloxygen 3,144 22 240 Hydrogenlox ygen 3,517 16 360 Hydrogedfluorine 4,756 10 390

The Russians use the kerosene/oxygen combination of fuels exclusive- ly in their launch boosters. These propellants were also used on the first stage of the Saturn V rocket used in the Apollo missions to the moon, but the United States regularly uses liquid hydrogen and oxygen for propel- lants in a majority of their launch vehicles including the main engines of the Space Shuttle. The difference in the use of these fuels is mostly due to their ease of handling; liquid kerosene (also known as RP-1) is much eas- Propulsion 61 ier to store and transport than liquid hydrogen. This is particularly demon- strated by the hydrogedfluorine combination which has been used in some special situations to take advantage of the higher specific impulse, but is not used regularly due to the highly toxic nature of fluorine.

Example Problem:

The space shuttle main engine (SSME) uses a liquid hydrogedoxy- gen propellant combination. Mass flow rate to each engine is 466.6 kg/sec and the combustion chamber pressure is 20.5 x lo6 N/m2. Determine the standard sea-level exhaust velocity, thrust, and specific impulse for a single SSME. (Assume pe = ps.l. = 1.01 x 16N/m2; y = 1.2; and g = 9.8 1 ds2)

Solution: Using the values for To, R, and M (converted to kg/mole) for the hydrogen/oxygen fuel combination given in Table 3- 1:

v, = 3,589.4 m/sec I,, = 365.9 sec T = 1.68 x lo6 N = 376,835 lbs thrust For comparison, a fully loaded space shuttle (including fuel tank and solid rocket motors) weighs over 4 million pounds, requiring three SSMEs and two solid rocket motors that deliver over one mil- lion pounds thrust each just to get off the launch pad.

Nozzle Design

If we look at a closed volume approach to the thrust equation, the situ- ation would look like that shown in Figure 3-3. In this case our control volume merely envelopes the rocket and we must consider forces which interact with, and through, this boundary. Summing the forces over the entire volume we find:

ZF = be+ (Pe - pm)& (3-7) The first term on the right-hand side of equation 3-7 represents the force due to the propellants passing through the control volume at velocity v,. This should be recognized as the same as the thrust term defined earlier. 62 Introduction to Space Sciences and Spacecraft Applications

P

P

P P Figure 3-3. Closed volume forces. An unbalanced force may exist on a rocket due to differences between the exhaust pmssure and ambient pressure.

The second term shows that the rocket may also feel a force on it due to a difference between the exhaust pressure and the ambient pressure in which the rocket is operating. This term is called the pressure thrust. The propellant exhaust pressure (like the exhaust velocity) is due to the aero- dynamic and thermodynamic expansion of the propellant through the con- verging-diverging nozzle, the design of which affects performance depending upon the particular operating environment. Three situations must be considered, as depicted in Figure 3-4.

Overexpansion. In this case, pe is less than p-, indicating that the exhaust gases have been overexpanded. This is undesirable because it produces a net pressure force opposite that of the rocket thrust (the p- in front of the rocket is not offset by the pe at the exhaust area).

Underexpansion. If the exhaust nozzle produces a situation where pe is greater than p-, the propellant gases will have been underexpanded This case looks, at first, to be desirable because the additional pressure force in the same direction as the rocket thrust should help push the rocket for- ward. However, we may recall that the thermodynamic equation for the exhaust velocity (eq. 3-4) includes a term with the ratio of exhaust pres- sure over combustion chamber pressure. We wished to minimize this term, which is why a high combustion pressure is desirable, but a high exhaust pressure increases this term. In fact, the loss of thrust due to a Propulsion 63

Overexpansion Underexpansion Perfect Expansion Figurn 3-4. Nozzle exhaust expansion. It is desirable to minimize the difference between exhaust pressure and ambient pressure. higher exhaust pressure in this equation far exceeds the gain in thrust due to the pressure thrust term.

Perfect Expansion. In this case, the exhaust nozzle expands the propel- lant gases such that pe is exactly equal to pD.and the pressure thrust term equals zero. By reason of default of the other cases, this is the best situa- tion when considering the pressure thrust term and nozzle design. The practicalities of nozzle design mean it is not always possible to maintain a perfect expansion condition throughout the operating range of the rocket. For instance, launch vehicles are used to carry payloads from the surface of the earth, through the atmosphere, into the near vacuum of space. Although variable exhaust area nozzles have been designed to con- tinually match exhaust pressures with the ambient conditions during tran- sit through the atmosphere, these systems are too complex to be practical and the thrust benefit is not worth the expense. In most cases, a median operating pressure is chosen and the nozzle designed to expand the pro- pellants to this pressure, accepting the small changes in thrust produced at other pressures. For systems that operate only in space, the situation is simpler and the nozzle is designed to expand the exhaust to as low a pres- sure as possible. This results in large, long rocket nozzles as exemplified by the Apollo service module nozzle depicted in Figure 3-5.

ORBIT ESTABLISHMENT AND ORBITAL MANEUVERS

As was mentioned at the end of Chapter 2, a spacecraft’s propulsion requirements may include delivery to space, maneuvering into position, and maintenance of the spacecraft position and orientation. While posi- tion-keeping systems are usually designed as an integral part of the space- 64 Introduction to Space Sciences and Spacecraft Applications

Figure 3-5. Apollo Service Module propulsion systems. The large aft nozzle was used for orbital maneuvers and the smaller nozzles around the body were used for . craft, rockets that deliver a spacecraft to its operating position are more likely to be existing systems which are chosen for specific performance characteristics and to which the spacecraft must be designed for compati- bility. Specific systems that are available to meet these requirements will be presented in Chapter 8. The following sections describe the propulsion aspects that must be taken into account when discussing these systems.

Orbit Establishment

Several factors are involved in reaching the desired orbit, including establishing the correct burn-out conditions, staging, launch timing, launch pad location, and launch direction. Figure 3-6 depicts a typical launch situation from the U.S. launch site at the Kennedy Space Center (KSC) in Florida.

Burn-out Conditions. The launch vehicle both accelerates and raises the payload from the surface of the earth along a predetermined launch path (trajectory). The launch path must intersect the desired orbit, and the inter- section must occur such that the rocket velocity (vector) and altitude cor- respond with those for the desired orbit ut that point. Since the rockets cease burning fuel at this point, the velocity and altitude achieved are known as the launch vehicle bum-out conditions. If the burn-out condi- tions do not match the orbit characteristics at that point, within some allowable tolerance, adjustments using supplemental propulsion systems may be required. Propulsion 65

STAGING

Figure 3-6.Orbit insertion. The launch vehicle must intersect the desired orbit at the proper altitude and with the proper velocity (magnitude and direction).

Staging. Multistage rockets are used to more efficiently (and practically) accelerate payloads to the high orbital velocities required. The separate stages are designed to optimize the burn-out velocity of each stage with respect to the overall mass the stage is accelerating. Upon burn-out, the empty stage is discarded and the next stage, optimized with respect to the remaining mass, takes over. The process is repeated until the desired velocity is achieved. A relatively simple relationship can be used to get an approximation of the number of stages required to deliver a payload to a particular orbit using a propulsion system of known characteristics:

mo - MR=--e (k) (3 - 8) *bo In equation 3-8, MR represents the mass ratio or ratio of the original mass (m,) to the mass at burn-out (mb) for each stage. Use of equation 3-8 assumes that this ratio is constant for all stages of the rocket. The velocity required to establish the orbit is represented by v,. This number will be higher than the orbital velocity of the desired orbit because it takes into account the excess thrust needed to operate through the atmosphere, 66 Introduction to Space Sciences and Spacecraft Applications elevate the payload to the orbital altitude, and redirect thrust during ascent to steer the rocket along the desired trajectory. The rocket exhaust veloc- ity, also assumed constant for all stages, is given by v,. Finally, n repre- sents the number of stages of the rocket.

Example Problem:

Although orbital velocity for a circular orbit at 300 km altitude is 7.73 km/sec, a typical velocity that a launch vehicle may be required to deliver to reach this orbit is 9.5 km/sec. Assuming use of space shuttle main engines for the propulsion system, determine the mass ratio for a single-stage-to-orbit (SSTO) rocket.

Solution:

Using the exhaust velocity found earlier for the SSME and remembering to use compatible units:

MR = 14.11

This indicates that the single stage would have to have about 14 times the mass of the payload in order to reach the given orbital velocity. This is actually a somewhat respectable mass ratio and illustrates the effectiveness of the SSME design. However, it must be remembered that equation 3-8 gives only a crude approximation due to the assumptions mentioned earlier. Actual stage optimization is a complicated, iterative process that takes into account the differ- ent thrust and mass ratio characteristics for each stage.

Launch Timing (Windows). In the preceding chapter we saw that the plane of an orbit is fixed inertially in space while the earth rotates beneath this plane. If it is desired to place the spacecraft into an orbital plane with a particular inertial orientation, the launch will have to be timed so as to occur just as the launch site rotates beneath the desired orbital plane, as depicted in Figure 3-6 for a launch from the Kennedy Space Center. In most instances, a plus or minus time period around the optimum time of launch is specified and is known as the launch window. If launch does not occur during this time period, the launch will have to be delayed until Propulsion 67 the earth rotates the launch site to another intersection with the desired orbital plane. For some launches, such as interplanetary missions, launch opportunities may be infrequent and launch windows may be as small as only a few minutes. Countdowns for launches are based on completion of all prelaunch activities before entering the launch window.

Launch Pad Location and Launch Direction (Azimuth). The direction in which the launch vehicle is fired from the launch pad is known as the launch azimuth. Azimuths start at 0' for a northerly launch and increase in a clockwise direction around the compass, as depicted in Figure 3-7. hunch pad location and launch azimuth are important, as they establish the inclination of the resulting orbit. This concept is most easily understood by looking at hypothetical launches from three different launch sites. Refer to Figure 3-7 and remember that the plane of the orbit cuts through the launch site and the center of the earth and contains the velocity vector of the spacecraft. Also remember that orbital inclinations are only described between 0' and 180" (refer back to Figure 2-6 in Chapter 2). Finally, the component of velocity imparted to our launch vehicle due to the rotation of the earth is initially ignored in this discussion but shall be reintroduced later.

Equatorial Launch Site. Launch sites located on the equator (Oo latitude) can attain any inclination orbit directly by simply launching in that direc-

0" azimuth

......

Figure 3-7. Launch azimuth versus orbit inclination. Launch site latitude and launch direction (azimuth) determine orbit inclination. 68 Introduction to Space Sciences and Spacecraft Applications tion from the pad. For example, a northerly launch (0" azimuth) would send its payload into an orbit passing over the poles continuously (ignor- ing the component of velocity imparted to our launch vehicle due to the rotation of the earth). This is the polar orbit described in the previous chap- ter with a 90" inclination. Note that a southerly launch (180" azimuth) also produces a polar orbit. In fact, we could achieve exactly the same (iner- tially-fixed) orbit as a northerly launch if we simply waited 12 hours for the launch site to rotate to the other side of the earth, and launched south! An easterly launch (90' azimuth) from this launch site would establish an orbit of 0" inclination-an equatorial orbit-circling the earth over the equa- tor constantly. An orbit with a component of its velocity (vector) in the same direction as the (easterly) rotation of the earth is known as aprograde orbit. Launching to the west (270" azimuth) would also establish an equatorial orbit, but the velocity would be opposite that of the rotation of the earth. An orbit with a component of its velocity vector opposite that of the rotation of the earth is called a retrograde orbit. The inclination of our retrograde equa- torial orbit would be 180", due to the fact that the angular momentum vector would point opposite that of the prograde equatorial orbit. Launching from our equatorial launch pad at a 45" azimuth results in an orbital inclination of 45". Note that launching at 135" azimuth also pro- duces a 45" inclination orbit. This is similar to establishing a polar orbit by launching north, or waiting 12 hours and launching south to enter the same inertially-fixed orbit. As you may now deduce, from our equatorial launch site we can enter any inclination orbit directly by simply launching in a compatible azimuth direction. You may also have noticed that launch azimuths between 0" and 180" result in prograde orbits with inclinations between 0" and 90". Launch azimuths between 180" and 360' produce retrograde orbits char- acterized by inclinations between 90" and 180".

Polar Launch Site. Next consider a launch site located on the North Pole (90" north latitude). The only direction toward which to launch is south (180" azimuth). This means that the only orbit that can be entered into directly from this launch site is a polar orbit of 90" inclination. Notice how the inclination corresponds to the launch site latitude.

Mid-latitude Launch Site. Now consider a launch site located at mid-lat- itudes, say 28.5"north. If we launch directly north (or directly south) from this point we would still establish a polar orbit of 90" inclination (again Propulsion @ ignoring the component of velocity imparted to our launch vehicle due to the rotation of the earth). However, if we launch directly east (90" azimuth), we find that the resulting orbit has an inclination of 28.5" cor- responding to the launch site latitude (refer to Figure 3-7). Launching a little northerly or even a little southerly (greater or less than 90" launch azimuth) from this location results in an orbital inclination greater than 28.5", meaning that 28.5" is the minimum inclination orbit into which we can launch directly from this site. The conclusion one should draw from this discussion is that the incli- nation into which a spacecraft can be launched directly is limited by the geographical location (latitude) of the launch site. If an orbit of lower inclination than the launch site latitude is desired, then an orbital adjust- ment (plane change) must be performed requiring additional fuel andor propulsion systems, which generally means increased weight and cost. The 28.5" latitude example used in the above discussion corresponds to the latitude of the Kennedy Space Center. Figure 3-8 shows the relation-

NORTH LATITUDE DEGREES

Flgure 3-8. United States launch sites. This figure shows the relationships between launch azimuths and orbital inclinations for U.S. launch sites. 70 Introduction to Space Sciences and Spacecraft Applications ship between launch azimuth and resulting orbital inclinations from both the Kennedy Space Center and Vandenberg Air Force Base in southern California, the two major U.S. launch sites. Though theoretically unlimited, the actual launch azimuths used by the United States are restricted to the directions that allow jettison (and in the case of the space shuttle, recovery) of the initial stages of the launch vehi- cle over the ocean. Vandenberg is used primarily to launch spacecraft into polar or the slightly retrograde sun-synchronous orbits, and, for the safe- ty reasons just stated, launch azimuths are always to the south.

Earth Rotation. In the previous discussions we ignored the rotation of the earth for simplicity in relating launch azimuths to orbital inclinations. However, the earth rotates at an appreciable rate which contributes a sig- nificant initial velocity to the launch vehicle. This initial velocity assists launches to the east, with the greatest assistance realized for 90" azimuth launches. This easterly component of velocity must be compensated for in order to launch into a truly polar orbit, and must be overcome in order to launch into a retrograde orbit. Launching in any azimuth other than 90" decreases the amount of payload a particular launch system can deliver to orbit compared to a direct easterly launch. This explains the popularity of 28.5" inclination orbits for U.S. space missions.

Orbit Adjustment

It was noted earlier that if the launch vehicle does not establish the desired orbital position and velocity precisely, or if the initial orbit is not the spacecraft's operating orbit, then an orbit adjustment will be required. We differentiate orbital adjustments from station keeping or attitude adjustments by the relatively larger Av requirements of orbital adjust- ments, as quantified in Chapter 2. Missions requiring orbital adjustments may incorporate completely separate propulsion systems to accomplish these maneuvers, while station keeping and attitude adjustment (if per- formed by a propulsion system) may be accomplished by a system inte- grated into the design of the spacecraft. Addressing all possible types of orbital adjustments and their associat- ed propulsion requirements is beyond the intent of this book. However, an excellent and commonly performed example exists which can be used to describe both the magnitude of adjustment required and the typical types of systems used to perform these maneuvers. Propulsion 71

Figure 3-9. Geostationary orbital transfer. The plane-change maneuver is performed out at the apogee of the transfer orbit to minimize the Av required.

Geostationary Transfer. Figure 3-9 illustrates the transfer of a satellite from a typical low-earth "parking" orbit of 28.5" inclination to a geosta- tionary orbit. The first maneuver establishes the spacecraft on a transfer orbit which is an ellipse that lies in the same plane as the parking orbit with its perigee at the same altitude as the parking orbit and its apogee at geostationary altitude. This should be recognized as a Hohmann transfer orbit described in Chapter 2. Note that in order to intersect the final desired equatorial geostationary orbit, the apogee (and, thus, the perigee) of the 28.5" inclined transfer orbit must be located over the equator. As the satellite approaches this point over the equator, a perigee kick motor (PKM)is fired to inject the spacecraft into the transfer orbit from the parking orbit. The change in velocity required at this point can be found from the relationships given in Chapter 2. At the transfer orbit apogee, two changes to the orbit must occur in order to establish a geostationary orbit. First, the elliptical transfer orbit must be circularized at the geostationary altitude. If this is done as a sep- arate maneuver within the same plane as the transfer orbit, the Av required would be the same as for a Hohmann transfer. However, to be geostation- ary, the orbit must also be equatorial, so the plane of our spacecraft orbit must be changed from 28.5" to 0". The Av relationship for a simple plane, given earlier in Chapter 2, is repeated here to illustrate an important point: Ai "simple plane change = 2v sin - (2 - 22) 2 This relationship reminds us that plane changes performed at low orbital velocities minimize the required Av. That is why, in geostationary 72 Introduction to Space Sciences and Spacecraft Applications transfers, the plane change is done out at geostationary altitude and not at the parking orbit altitude. In actual geostationary transfers, an apogee kick motor (AKM) is used to perform both the circularization of the transfer orbit and the plane change simultaneously, which results in a savings in Av required over performing separate maneuvers. PKMs and AKMs may be either liquid or solid fueled. Because only a single bum of known Av is required at the "kick" points of a geostation- ary transfer, solid motors provide a simple way of performing this func- tion and are commonly used. However, if higher performance or a restart capability is required, a liquid-fueled motor may be more suitable. Dif- ferent PKMs and AKMs as well as currently available orbital maneuver- ing systems and launch vehicles are presented in Chapter 8. Before leaving this subject, it is worth noting that the former Soviet Union's major launch site, the in Khazakstaan near the city of Tyuratam southeast of Moscow, is at 45.6" N latitude, as shown in Figure 3-10. Due to range safety reasons and the proximity of China to the east, allowable launch azimuths result in a minimum orbital inclination of 51" from Tyuratam. It should come as no surprise that the Russian Mir space station is in a 51" inclination orbit. Also, as with the Apollo-Soyuz mission, future joint U.S.-Russian missions, including the international space station, will probably be conducted in 5 1O inclination orbits, as the penalty paid by the United States to launch into a higher than 28.5" inclination orbit is far less than that which would be required by the Russians to change the plane of their orbit from 51" to 28.5". As a last note, the French Ariane launch vehicle has become a major launcher of satellites into geostationary orbit. Much of this success is due to the fact that the Ariane is launched from the Guiana Space Center near Kourou, located at 5.2" N latitude on the northeast coast of South America. Since only a small plane change is required to establish a geostationary orbit from this site, more of the payload launched can go into spacecraft weight and less to propulsion systems making the Ariane attractive to many com- mercial geostationary satellite providers. Figure 3- 10 shows this launch site along with the locations and latitudes of some of the world's other current major orbital launch sites.

Station-keeping and Attitude Control

In Chapter 2 we ignored the many sources of perturbations that con- stantly act on the orbit and orientation of a spacecraft. If it is important Propulsion 73

1 I Figure 3-10. Launch sites. Launch site latitude determines the minimum inclination that a satellite may be launched directly into.

that a spacecraft maintain a specific orbit and/or attitude with respect to the earth or stars, then additional propulsion systems may be required. Compared to launch and orbital transfer vehicles, station-keeping and atti- tude control systems are made up of much smaller thrusters which mea- sure their forces in pounds (or Newtons) or fractions thereof. An orbital transfer vehicle may be capable of delivering a Av of several kilometers per second in order to make major changes to a spacecraft orbit, but a sta- tion-keeping propulsion system is designed to make fine adjustments to an orbit, delivering Av’s measured in meters per second. Attitude control thrusters are even smaller, designed to simply rotate the spacecraft body and not affect the orbit significantly. As mentioned earlier, these systems are usually custom-designed as an integral part of a spacecraft, and the thrusters must be positioned about the spacecraft body at precise locations to provide the correct torques and magnitudes. If you refer back to Figure 3-5 you will notice the groups of small thrusters placed around the body of the Apollo service module. Associated with these thrusters, propellant tanks must be included to store the required amount of fuel for the mission lifetime. In many cases, the amount of onboard fuel becomes the limiting factor in useful space- craft lifetime, and more than one perfectly-operating spacecraft has become useless due to fuel depletion. The propellant used must be capa- ble of on-orbit storage. If a liquid propellant is used, the spacecraft may 14 Zntroduction to Space Sciences and Spacecraft Applications have to maintain the tanks and lines within acceptable temperatures so the fuel does not freeze and burst its containers. The loss of the $1 billion Mars Observer as it was approaching its destination in August 1993 was attributed to burst propulsion system pipes. Also, since the exhaust gases may remain in the Vicinity of the space- craft, the propellant must not become a source of contamination to the spacecraft systems or the operation of the payload. Pressurized gases such as nitrogen, which are less of a source of contamination than many other propellants, are often used for attitude control; however, their perfor- mance (specific impulse) is less than that of liquid propellants and may not be acceptable for higher thrust purposes. A commonly used liquid mono-propellant for station-keeping and attitude control is hydrazine (N2H4), of which there are several variations (monomethylhydrazine [MMH] and unsymmetrical dimethylhydrazine [UDMH] are common examples). Hydrazine has many desirable characteristics in terms of stor- age and contamination, and delivers a very acceptable specific impulse of over 200 seconds.

REFERENCES/ADDITIONAL READING

Anderson, J., Zntroduction to Flight, 2nd ed. New York: McGraw-Hill, Inc., 1985. Sutton, G., Rocket Propulsion Elements, 5th ed. New York: John Wiley & Sons, Inc., 1986. C. Cochran, D. Gorman, and J. Dumoulin (Eds.), Space Handbook. Alabama: Air University Press, 1985.

EXERCISES

1. Determine the theoretical exhaust velocity for the liquid hydrogedliq- uid oxygen fueVoxidizer combination given in the reading. Consider the ratio of exhaust pressure to combustion chamber pressure (pJp,) and the ratio of specific heats (y) to be 0.001 and 1.4, respectively. 2. Recompute the above exhaust velocity with an increase and then a decrease in the pressure ratio of a factor of ten (Le., p$po = 0.01 and pJp, = 0.0001). Propulsion 75

3. Compare the answer obtained in Exercise 1 to the exhaust velocities obtained for the other fueyoxidizer combinations given in the reading. Use the same values for pJp, and y given in Exercise 1 above. 4. Compute the specific impulses for the three fueyoxidizer combina- tions (based on the answers from Exercises 1 and 3 above) and com- pare them to those given in the reading. What might be the reason for any differences? 5. The Saturn V F-1 first-stage booster rocket used a kerosene/oxygen propellant combination. Mass flow rate to each engine (there were five in the first stage) was 5,736 Ib/sec and the combustion chamber pres- sure was 1,122 psi. Determine the standard sea level exhaust velocity, thrust, and specific impulse for one of the engines (assume y = 1.2). What was the total sea level thrust of the Saturn V? Give your answers in both English and metric units. 6. It is desired to launch a with a ratio of original mass to burn-out mass for each stage of 2.0. For a nozzle design which main- tains the exhaust velocity at 3,000 dsec, determine how many stages would be necessary to place a payload into a circular low-earth orbit at an altitude of 200 km, with a trajectory resulting in VJV- = 1.1. 7. Compute the three separate Av's required for the geostationary transfer described in the reading. Assume an initial circular parking orbit of 450 km altitude at 28.5" inclination. CHAPTER 4 Spacecraft Environment

Spacecraft operate in an environment totally different from that experi- enced by systems on earth. In some cases, this environment is relatively benign compared to what earth-based systems must withstand (hurricanes, for example), but in many ways the space environment is harsh on the equipment (and people) sent out to perform there. The space systems engineer must understand this environment and design spacecraft to oper- ate within it. In addition, the spacecraft must be capable of performing its mission through this environment. The effects of the environment on typ- ical spacecraft missions will be discussed in subsequent chapters.

THE SUN

The sun is the major environmental influence on the earth and the space around it. This section presents a brief description of the sun and discuss- es the radiations it emits into space. Following sections describe the prop- erties of these radiations in the vicinity of earth and the environment that results from the interaction of the two. Finally, the effects of this environ- ment on the operation of systems in space is discussed.

Cosmology

Cosmology is the study of the evolution of the universe and includes the theory of how our sun came into being. It is believed that 12 to 18 billion years ago the universe had a violent beginning in an expanding fireball known as the “Big Bang.” The Milky Way galaxy represents one of the countless lumps of the products of that explosion, made up mainly of hydrogen and helium. Our solar system began as a cloud of interstellar gas (also mostly hydrogen) which probably contained the debris of the destruc-

76 Spacecraft Environment 77 tion of previous generations of stars along with the primordial elements of the creation of the universe. It is theorized that some sort of shock wave, perhaps from a nearby stellar explosion (the fate of all stars) or a galactic wave front associated with the movement of the spiral arms of our galaxy, disturbed the cloud of gas, causing it to condense in places. These areas of increased density and gravitational attraction pulled in nearby molecules, and the gas cloud began to spiral inward forming a disk-like shape. Some local clumps of material formed away from the main core and, through accretion (collision and fusing) with other clumps, eventually coalesced into what are now the planets and moons. The ring of asteroids around the sun between the orbits of Mars and Jupiter are thought to be a region where this process did not continue long enough to form another planet. The planets all basically lie in the plane of the sun's equator, known as the ecliptic, and travel in the same direction around the sun. As the mass at the center of this activity grew larger, pressure and tem- perature began to rise as the elements packed in tighter and tighter until conditions reached a critical point and the core burst into spontaneous nuclear fusion. Many more processes occurred in the 4.6 billion years or so it has taken the solar system to stabilize into what we see today, but at that moment a new star was born. Our sun is a typical star in a typical galaxy containing an estimated 200 billion stars in a universe of perhaps more than a billion galaxies. The sun is located about 30,000 light-years (1 1.y. = the distance light travels in one solar year = 9.5 x lo1*km) from the center of our spiral galaxy which is about 100,OOO 1.y. across. The sun has an orbital period of about 200 million years and an orbital velocity of about 250 km/sec around the Milky Way.

Structure of the Sun

The sun has a mass (M,) of 2 x 1030 kg composed of about 78% hydro- gen, 20% helium, and 2% heavier elements by weight. These materials exist mainly in the form of a plasma, a homogeneous mixture of ionized elements and their dissociated electrons. Solar density varies from around lo5 kg/m3 at the core (five times that of uranium) to lo4 kg/m3 at what we define as its surface (the photosphere). Solar radius (R) at the photo- sphere is around 696,000 km. For comparison, the density of water is 1,000 kg/m3, the mass of the earth (Me) is 6 x kg, and the radius of the earth (%) is 6,378 km. The sun has a rotation rate which differs with solar latitudes-approximately 25 days per rotation at the equator and up 78 Introduction to Space Sciences and Spacecraft Applications

RB - 1 solar radius

Figure 4-1. Solar regions. The sun’s characteristics change with distance from its center.

to 30 days per rotation in the vicinity of the solar poles. The sun is com- posed of six major regions as shown in Figure 4-1.

Core. The core is the region of the sun in which the conversion of mass into energy occurs. Temperature in the core is about 15 x lo6 OK and the density is about 110,000 kg/m3, conditions sufficient to cause the fusion of hydrogen nuclei into helium nuclei with an associated release of ener- gy. Approximately 4 billion kilograms of mass is converted into energy every second, producing about 4 x kW of power. There is enough fuel in the sun for this process to continue for at least a few billion more years, so there’s no need to panic just yet.

Radiative Zone. In the region of the sun known as the radiative zone, the energy produced in the core is transported slowly upwards by means of absorption and reradiation of gamma rays, X-rays, and ultraviolet (UV) photons from layer to layer.

Convective Zone. Through the convective zone, the major transportation of energy is the convection (carrying) of energy by bubbles of hot gas that boil their way up to the surface. Through a telescope, these bubbles can be seen as a textured pattern on the sun called granulation. It is estimated that the transport of the energy created in the core takes about 10 million years to make its way through the radiation and convection zones to the photosphere, where it is released to the solar system. Spacecraft Environment 79

Photosphere. The photosphere is a relatively thin layer (only about 300 km thick) of hot gas which absorbs the energy produced beneath it and then reradiates this energy into space. All electromagnetic radiation, including the visible wavelengths we see from earth, emanates from this region of the sun. Since this is what we perceive as its shape visually, this layer is defined as the “surface” of the sun even though the solar body extends far beyond. The temperature of the photosphere is roughly 6,000 OK which, as we will see, determines the characteristics of part of the solar radiation.

Chromosphere. The layer of gas above the photosphere is known as the chromosphere and corresponds roughly to the “atmosphere” of the sun. A temperature profile exists through this layer which initially decreases to about 4,000 OKthen increases rapidly to about lo6 OK (although temper- ature loses its familiar meaning with such large distances between parti- cles). The density within the chromosphere is sufficiently low that the electromagnetic radiation from the photosphere is not appreciably affect- ed in its outward passage through this region.

Corona. Finally, the corona is a layer of gases that extends millions of miles away from the sun. Temperatures remain high, and the density of the gases is on the order of lo-” kg/m3. The corona can be seen as a streaming pattern around the edges of the moon during a full solar eclipse.

Solar Radiation

The environment in the vicinity of the earth is affected by two major products of the sun’s nuclear furnace. The most familiar is the electro- magnetic radiation of which visible light is a part. The second is the out- pouring of high energy solar particles.

Electromagnetic Radiation. Electromagnetic radiation is energy which propagates at the speed of light in a wave-like manner. Figure 4-2 depicts a typical sine-wave pattern used to describe an electromagnetic wave and illustrates some basic characteristics with which we will be dealing. Wavelength (A)represents the distance (meters, microns) between suc- cessive peaks of a particular wave pattern. Since electromagnetic energy travels at the speed of light, the number of peaks that pass a fixed point in a second of time is known as thefrequency (f) of the radiation. The rela- 80 Zntroduction to Space Sciences and Spacecraft Applications

wavelength

Figure 4-2. Electromagnetic wave. A simple sine-wave depicts some of the characteristics of an electromagnetic wave. tionship between frequency and wavelength is given in equation 4-1 in which c represents the speed of light (3 x lo8 dsec in a vacuum).

C f = - cycles per sec ond (Hz) (4-1) h Two other characteristics of electromagnetic radiation, which we will look at more closely later, are the amplitude and the phase. Amplitude is represented by the height of the peaks of the sine-wave, and phase describes a particular location along the wave. Figure 4-3 shows many of the regions of the electromagnetic spectrum with which we will be dealing. Notice how small the range of visible wavelengths is. Remember that as the frequency increases, the wave- length decreases, as explained by equation 4-1.

Blackbody Radiation. A blackbody is a theoretical body which has the properties of being a perfect radiator (and absorber) of energy. As we will see in the next few sections, the characteristics of the energy radiated by a blackbody is a function of the body’s temperature. Though it is a theo- retical concept, many bodies resemble blackbodies, at least closely or over a particular range of wavelengths, and can be approximated by blackbody relationships.

Stefan-Boltvnann Law. In 1879 an Austrian physicist, Josef Stefan, with help from his contemporary Ludwig Boltzmann, demonstrated that the total energy radiated by a blackbody increased as a function of the fourth Spacecraft Environment 81

ULTRA- INFRA- VIOLET RED

3l’Hz 314Hr 3000Hz 300MHz 300KHz frequency I I I I I I I I I I I I I I 10-12m 1 nm lpm lmm lm dm 16, wavelength Figure 4-3. Electromagnetic spectrum. The major categories of electromagnetic radiations are indicated. power of the body’s absolute temperature. In terms of power output per unit area, the Stefan-Boltzmann relationship is:

where the temperature T is in OKand 0 represents the Stefan-Boltzmann constant (o = 5.76 x 1W8W/m2 OK4). The temperature of the sun’s photo- sphere is approximately 6,000 OKwhich, when substituted into the above relationship, results in a value E, = 74.7 x IO6 W/m2 indicating that each square meter of the sun’s surface radiates more than 70 megawatts!

Solar Constant. We can estimate the total power output of the sun by mul- tiplying the number just obtained from the Stefan-Boltzmann relationship by the surface area of the sun. Using the radius of the sun at the photosphere (R, = 696,000 km), this total energy (power) can be found from:

P, = E, x 4nR; (W) (4-3)

Assuming negligible energy losses in space, the same amount of total energy must pass through any two spheres drawn around the sun. The earth circles the sun at an average distance of 149.5 x IO6 km (represent- ed by re, this has been adopted as a unit of distance measurement known as one astronomical unit or 1 A.U.). Using equation 4-3 to equate the energy emitted by the solar surface to that passing through a sphere of 1 A.U. radius, we get: 82 Introduction to Space Sciences and Spacecraft Applications

E, x 4zR: = E, x 4m-Z (4-4) where E, represents the solar-generated energy per unit area in the vicin- ity of the earth. Of course, due to the increased surface area of the larger sphere, the energy in the vicinity of the earth is less than that at the sur- face of the sun. Using the value obtained for a 6,000 OK sun, we get a value of E, = 1,631 W/m2. This is the amount of energy that would impinge on a square meter area at a distance of 1 A.U. due to the energy output of the sun and is an important value when considering energy pro- duction via solar cells or for determining thermal inputs for a spacecraft. This value (E,) is known as the solar constant. Even though recent studies have shown that solar power output may fluctuate slightly (which may have affected the climate on earth severely at times in the past, per- haps contributing to the ice ages), for our purposes we shall consider it a constant value.

Wien’s Displacement Law. In 1895, German physicist Wilhelm Wien, discovered that the wavelength corresponding to the maximum energy output for a blackbody at a particular temperature could be found from a simple relationship:

For a temperature of 6,000 OK, equation 4-5 yields a maximum-energy radiated wavelength of 0.483 pm which corresponds to the yellow-green light frequencies close to the middle of the visible spectrum. The fact that human vision has adapted to take advantage of the maximum portion of the solar energy output seems like a good argument in support of the the- ory of evolution.

Planck’s Law. In 1899, another German physicist, Max Planck, derived a relationship that combined the findings of his predecessors and described the distribution of blackbody radiation as a function of temperature and wavelength: Spacecraft Environment 03 where K = Boltzman’s constant, h = Planck‘s constant, and the other terms are as described previously. If this relationship is plotted for a temperature of 6,000 OK, the result would be as shown in Figure 4-4.The figure also shows the relationship plotted for a body at a temperature of 300 OK, which is the average tem-

104

102

1

io-.

106

.oi .I 1 io 100 io3 io4

WAVELENGTH (pm) Figure 4-4. Planck’s Law. The wavelength distribution of energy emitted by a body depends on its temperature. perature of the surface of the earth. Both the earth and the sun behave as blackbody emitters over much of the electromagnetic spectrum (as we will see later). The sun emits energy at essentially all wavelengths of the spectrum with the maximum energy output in the visible frequencies as we determined earlier. The energy levels for the earth are significantly lower (Figure 4-4is plotted using logarithmic scales) and the peak of the curve is shifted to a higher wavelength (lower frequency). Due to its lower temperature, the earth radiates its maximum energies in the infrared wavelengths, as may be determined from equation 4-5. One way to picture this temperature/wavelength relationship is to imagine a metal rod heated in a furnace. As its temperature is increased, the rod starts to glow, getting red then white hot, illustrating its radiation of different wavelengths with changing temperature. 84 Introduction to Space Sciences and Spacecraft Applications

Example Problem:

The sun’s actual average “surface” temperature is about 5,750 OK. Based on this temperature, determine the blackbody power output of the surface of the sun, the expected solar constant in the vicinity of the earth, and the wavelength and associated frequency of maximum solar energy output.

Solutions:

From equation 4-2: E, = 62.96 x lo6 W/m2

Using equation 4-4: E, = 1,364.7 W/m2

From equation 4-5: A,, = 0.50 pm

Finally equation 4- 1: f,, = 6 x 1014 Hz

Solar Magnetic Field. The dynamic nature of the sun, causing motions of the solar plasma, generates strong magnetic fields. Magnetic field strength is commonly measured in units of Gauss or gamma (1 G = lo5 y) The strength and direction of the magnetic fields change over time and with changing processes on the surface, which will be described shortly. How- ever, on average, the sun’s surface magnetic field strength is about 15 G, approximately 30 times the surface magnetic field strength of the earth (0.5 G at mid-latitudes).

Particle Radiation. In addition to the electromagnetic energy released, there is a significant amount of particulate matter emitted into space by the sun of which there are basically two major sources: solar wind and solar flares.

Solar Wind. The solar wind, actually the extension of the corona, is a steady stream of plasma particles given sufficient energy to escape the sun’s gravitational attraction and propagate into space in all directions. Through a process which implies that the product of the density and velocity of these particles remains constant with increasing distance from Spacecraft Environment 85 the sun, the velocity of this stream increases as it leaves the sun (to a point, after which it remains relatively constant).* As mentioned earlier, the sun rotates while generating a substantial mag- netic field. A characteristic of a plasma passing through a magnetic field is that the field lines become trapped and move with the plasma. As the solar wind propagates away from the sun, the solar magnetic field is dragged along but, due to its magnetic properties and the rotation of the sun, also affects the propagation of the plasma. The result is that the solar wind can- not be described as propagating in a straight path away from the sun. Actu- al propagation is three-dimensional and complex, making it difficult to predict the actual environment that may appear in the vicinity of the earth. On average, the velocity of the solar wind in the vicinity of the earth is about 500 km/sec (supersonic in the definition of the term) and the densi- ty is about five particles/cm3. Solar magnetic field strength is about 5 y, approximately 1/10,000* the surface magnetic field strength of the earth. As a final note, it is suspected that a boundary exists around 100 A.U. away from the sun where the streaming solar wind encounters the interstel- lar space environment defining a region surrounding the sun known as the heZiosphere.t Some futuristic propulsion schemes propose to use the solar wind, like a sailboat uses atmospheric winds, to travel between the planets.

Sunspots/Sohr Flares. The second major source of particle radiation is due to disturbances on the solar surface. The processes which produce sunspots and solar flares are intimately linked and must be described together.

Sunspots. Galileo, using the crude telescopes of his own design, was the first to notice “blemishes” or spots on the surface of the sun. Even though he had no idea of the processes that may have accounted for these spots, he did keep a close enough watch on them to determine that the sun

*The Ulysses spacecraft, a joint NASABSA mission launched from the Space Shuttle in October 1990, used a gravity assist from Jupiter to change its orbital plane out of the ecliptic (the sun’s equa- torial plane within which the planets more or less lie) to sample the interplanetary space environ- ment over the sun’s poles. The results show a difference in the velocity/density relationship from the ecliptic, which has yet to be fully explained. ___ - __ +Pioneer 10 (launched in December, 1973) and Voyager I and Voyager 2 (launched in September 1977 and August 1977, respectively) interplanetary spacecraft are on trajectories that will carry them beyond the solar system. In 1996, Voyager 2 will be around 45 A.U. from the sun, and Pioneer 10 and Voyager I will both be approximately 60 A.U., all traveling in different directions. These still- operating spacecraft continue to be monitored for information to help determine the extent and char- acteristics of the heliosphere. 86 Introduction to Space Sciences and Spacecraft Applications

rotates on its axis, which we now know is a rotation rate which differs with solar latitudes. As it turns out, this differing rotation rate represents part of the process which produces sunspots and solar flares. As mentioned earlier, magnetic fields passing through a plasma become trapped and move with the plasma. The differing rotation rate of the sun causes these trapped magnetic field lines to “wind up” and increase their field strength to the point where they produce a magnetic pressure on the plasma. In these areas, the gas pressure is lower, which lowers the gas temperature. As we have seen, the sun “looks” yellow due to the temper- ature at which its surface is radiating. Sunspots are darker-looking patch- es of lower temperature due to these areas of increased magnetic field strength at the surface. Magnetic field strengths of 1,500 G in these loca- tions are not uncommon.

Solar Flares. The actual cause of solar flares is still not completely under- stood. One explanation states that the magnetic field lines continue to wind up to a strength where they pop out from the surface of the sun, car- rying solar plasma into space. The escaping material may be accelerated by the field above the surface to velocities approaching relativistic speeds up to 106 dsec. Intense electromagnetic energy in ail wavelengths, from X-rays through radio, is also released by solar flares. As we shall see shortly, arrival of these high-energy particles and intense radiation to the vicinity of the earth creates a situation which can affect the operation of ground systems as well as spacecraft.

Solar Cycle. Records have been kept on the number of sunspots visible on the sun’s surface since the mid-l700s, and review of these records indi- cates that the number of sunspots changes over time with some regulari- ty. At times, no sunspots at all may be visible, while at other times over 100 sunspots may exist simultaneously. If one describes the period of lit- tle or no sunspot activity as solar minimum and the period of highest activity as solar maximum, the average time from one minimum or max- imum period to the next can be defined as the solar cycle. Though the solar cycle varies between approximately seven to thirteen years, the aver- age cycle has been found to be about eleven years. Also, on average, it seems to take less time to rise to solar maximum, around four years, than to settle back to solar minimum, around seven years. The next solar max- imum is expected to occur around the year 2000. Spacecraft Environment W

Sunspots seem to occur only around mid-solar latitudes, between around 60” north and south of the equator, forming closer to the equator as the cycle approaches minimum. A reversal of the magnetic polarity of the sun, from positive to negative lines of magnetic flux emanating from the geographical poles, occurs over each solar cycle. This variability in intensity, location, and magnetic polarity over time manifests itself in a changing and difficult-to-predict environment in the vicinity of the earth.

THE EARTH

The environment on and around the earth is affected in many ways by the electromagnetic and particulate radiations from the sun. Spacecraft must operate in and through the resulting environment, which includes the earth itself and its atmosphere as well as the space around it. Consequent- ly, the following sections describe some of the properties of the earth as well as the interaction of the earth with the solar radiations.

Electromagnetic Interaction

We previously determined the amount of electromagnetic energy that reaches the vicinity of the earth when we calculated the solar constant. This energy retains the blackbody spectral characteristics it had when it left the sun, the only difference being that the energy levels over the range of wavelengths has decreased. As this energy encounters the earth, the inter- actions affect both the earth and the energy itself. In some cases, these interactions are useful to the space systems user. For example, remote sens- ing techniques measure the absorption or transmission of solar radiation through the atmosphere to infer some of the characteristics of the atmos- phere from space, and most spacecraft use solar panels to convert some of this radiation to electricity to power the spacecraft systems. In other cases, solar radiations produce undesired effects. For instance, spacecraft may absorb excess infrared frequencies, necessitating the use of thermal control devices, and the ionosphere, described next, may introduce errors in trans- mitted signals. Spacecraft systems, including power generation and ther- mal control, and spacecraft applications, such as remote sensing and com- munications, are described more thoroughly later in the text.

The Ionosphere. The short wavelength (X-ray, UV) radiations from the sun have enough energy to “knock” electrons out of some of the constituents of 88 Introduction to Space Sciences and Spacecraft Applications

the atmosphere at high altitudes. This process is called photoionization and results in an electrically charged (ionized) region known as the ionosphere. This region begins somewhere between 50 and 70 km altitude and extends far into space. The ionosphere is characterized by the changes in the free electron concentration levels with altitude as shown in Figure 4-5. Figure 4-5 also shows the variability of the ionosphere. Solid lines indi- cate levels during times of high solar activity and dashed lines are associ- ated with low solar activity levels. Notice also the considerable difference between day and night characteristics. This is due to the cessation of the photoionization process when out of direct exposure to the sun. During these times the electron concentration is reduced due to the recombination of ions and electrons. This process proceeds slowly enough to allow a concentration of free ions and electrons to persist throughout the night. Some easily identifiable regions within the ionosphere can be seen in Fig- ure 4-5 associated with changes in the electron concentration with altitude. These regions (layers) are identified with letters (D, E, Fl, F2, or just F). This layering of the ionosphere is associated with the changes in atmos- pheric constituents with altitude which have differing ionization properties. We will return to our discussion of the ionosphere later when we discuss communications, remote sensing, and navigation satellites.

HIGH ACTIVITY

800 800

400

W n 2 200 !i a 150 100 80

60 I 1 I IO lo2 lo3 io4 io5 lo6

ELECTRON CONCENTRATION (cm3) Figure 4-5. The ionosphere. Electron concentration in the ionosphere changes with altitude, time of day, and solar activity. Spacecraft Environment 89

The Atmosphere. A chapter on space environment may seem like a strange place to include a discussion of the earth’s atmosphere, but the atmosphere is a factor in the operation of spacecraft in many ways. The following sections describe the atmosphere, along with its interaction with the sun’s electromagnetic radiations.

Composition of the Atmosphere. It is believed that the atmosphere was created as a result of the exhausting of materials from the interior of the earth by volcanic activity. Nitrogen, being relatively inert, remains as the most common atmospheric element. Diatomic oxygen was mainly pro- duced via the photosynthesis of water and carbon dioxide by plant life over millions of years. Argon and other elements make up the remainder of the earth’s atmosphere as shown in Table 4- 1.

Table 4-1 Composition of the Earth’s Atmosphere (blow 100 km)

Constituent Content (fraction of total molecules) Nitrogen (N2) 0.7808 (75.51% by mass) Oxygen (02) 0.2095 (23.14% by mass) Argon 0.0093 (1.28% by mass) Water vapor (H20) 60.04 (variable) Carbon dioxide (C02) 325 parts per million Neon (Ne) 18 parts per million Helium (He) 5 parts per million Krypton 1 part per million Hydrogen (H) 0.5 parts per million Ozone (03) 0-12 uarts Der million

Higher in the atmosphere, monatomic oxygen (single oxygen atoms) are created through photodissociation (splitting) of molecular oxygen by the sun’s radiation. Monatomic oxygen is highly reactive, and some of these atoms combine with diatomic oxygen to produce the ozone (0,)found in the atmosphere. Ozone is particularly important to life on earth because it absorbs much of the harmful ultraviolet radiation emitted by the sun.

Temperature Variation. The variation of temperature with altitude up to about 700 km through the atmosphere is shown in Figure 4-6. These val- ues are for a “standard” atmosphere which represents typical conditions at 90 Introduction to Space Sciences and Spacecraft Applications

thermosphere 700

600

500

50 - 400

300

200

100

160 180 200 220 240 260 280 500 1000 1500 2000

TEMPERATURE (OK)

Figure 4-6. The “standard” atmosphere. The temperature of the atmosphere varies with altitude.

mid-latitudes. This temperature variation delineates several distinct regions also identified in the figure.

Troposphere. The troposphere is an area of unstable air where most of the weather we experience is formed. The region is unstable due to the turbu- lent convection of air heated at the surface by the infrared energy that the earth emits due to its temperature, which is a result of the earth’s absorp- tion of solar radiation. As can be seen, the temperature decreases linearly through this region until we reach the tropopause-the transition region between the troposhere and the stratosphere-where the temperature remains relatively constant (isothermal).

Stratosphere. Unlike the troposphere, the stratosphere is relatively non- turbulent. Temperature increases with altitude through this region due to the continued absorption of the earth’s infrared radiation by H20 and C02 molecules and absorption of solar ultraviolet radiation by the ozone that exists at the top of this region.

Mesosphere. Above the isothermal layer at the top of the stratosphere (the stratopause), the mesosphere is a region where the temperature once again Spacecraft Environment 91 decreases with altitude. This is due to the reradiation of energy absorbed from the layers below which escapes into space.

Thermosphere. After the mesopause (another isothermal layer), the ther- mosphere represents a region in which the temperature increases rapidly with altitude. This is due, once again, to the direct molecular absorption of solar radiation by the atmospheric constituents at these altitudes.

Exosphere. In the exosphere, temperature remains relatively constant with altitude due to the low densities and the large mean free path (distance) that exists between molecules. In this region, energetic molecules may actually escape from the earth’s gravitational pull and be lost into space.

Density Variation. Decrease in the density of the atmosphere with increasing altitude is a result of the balance between the gravitational force on molecules of different masses and the thermal energy of these molecules. Though relationships have been developed that relate the change in density with altitude based on the standard temperature profile, they do not work well at the higher altitudes at which spacecraft operate. A good rule of thumb is that up to 100 miles altitude, the density decreas- es by a factor of 10 every 10 miles. Above this, the decrease becomes exponential as the lighter elements such as hydrogen and helium become more predominant. The atmosphere extends for thousands of miles above the earth’s surface, but in ever-decreasing densities. The density of the atmosphere at any particular altitude also varies with time due to the variability of solar emissions over the solar cycle and dur- ing solar flares. Increased emissions may heat up the atmosphere causing an increase in density at satellite altitudes which increases drag and may affect orbital lifetimes.

Electromagnetic Propagation Through the Atmosphere. As the previous discussions have indicated, the atmosphere absorbs much of the electro- magnetic energy radiated from both the sun above and the earth below. It also reradiates much of this energy in both directions as well, but the aver- age rates of absorption and emission are about equal, which results in a relatively constant average temperature around the globe. The mechanisms behind absorption and emission are rooted in chem- istry, physics, and quantum theory which show that these properties hap- pen at relatively distinct frequencies for different atoms and molecules. 9 Intmduction to Space Sciences and Spacecraft Applications

I radiation curve for 5900 OK black body

.solar radiation outside atmosphere

.2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0

WAVELENGTH (m) Figure 4-7. Electromagnetic propagation. Much of the solar radiation inci- dent on the earth is absorbed by the atmosphere.

Figure 4-7 shows the effects of this compatibility of certain elements with certain frequencies (wavelengths) over the range of radiations incident on the earth from the sun. The topmost curves in Figure 4-7 represent the spectral energies that arrive from the sun at the top of the earth’s atmosphere. The dashed line indicates the theoretical energies given by the relationships discussed ear- lier in this chapter for an idealized blackbody. As you can see, the corre- lation of this theoretical curve with the measured irradiation outside the atmosphere is quite good. The lower solid curve indicates the amount of energy, by wavelength, that reaches the earth’s surface. Although there is attenuation in most of the incident radiations, you can see that there are some wavelengths which are drastically (and some completely) attenuat- ed in their passage through the atmosphere. The associated elements which absorb these energies are also indicated on the figure.

The Greenhouse Effect. Figure 4-7 shows that much of the energy of the sun’s radiation does reach the earth’s surface. This incident radiation warms the surface, which then reradiates this energy-due to this temper- ature and Wien’s law-in the infrared frequencies. However, the C02and H20molecules in the atmosphere are highly absorbent in the infrared and, as a consequence, the atmosphere is heated, as was indicated in the earli- Spacecraft Environment 93 er descriptions of the atmospheric regions. As it stands, there is a thermal balance between the amount of energy absorbed and the amount reradiat- ed into space, but the above discussion shows the concern over increasing the carbon dioxide levels in the atmosphere as can happen through the large-scale burning of fossil fuels and forests.*

Particle Interactions

To discuss the characteristics of the solar particle radiations in the vicinity of the earth, we must first discuss the earth’s own magnetic field and the interaction of charged particles with this field.

Geomagnetic Field. As any good knows, the earth has a magnetic field on which the pointer of a compass can align and direction be deter- mined. One explanation for its existence is given by the Dynamo Theory which states that the field is generated by a complex and nonuniform rota- tion of the earth’s molten metallic core. As shown in Figure 4-8,the geo-

Figure 4-8. Geomagnetic field. The earth has a magnetic field that can be modeled as a huge magnet located near the earth’s center.

*It is believed that high atmospheric COz concentrations are a major contributor to the 700°F condi- tions on the surface of Venus-temperatures hot enough to melt lead! 94 Introduction to Space Sciences and Spacecraft Applications magnetic field can be modeled by a (quite large) dipole magnet placed within the earth and tilted about 11.5' from the earth's axis of rotation. This theoretical magnet would not reside exactly at the geometric center of the planet but would be displaced about 436 km off-center in the direc- tion of the Pacific Ocean.

Interaction of the Solar Wind and the Geomagnetic Field. When the solar wind encounters the Earth's magnetic field, an electromagnetic inter- action occurs that deforms the geomagnetic field as depicted in Figure 4-9. As the supersonic charged particles of the solar wind encounter the mag- netic field of the earth, a shock front is formed across which the solar parti- cles are deflected and slowed, similar to that which occurs in aerodynam- ics. The distance to the earth of the shock front is about 14 earth radii (14 Re) at its closest in the direction of the incoming solar wind, increasing greatly as the angle away from this direction increases. Inward from the shock front, the particles exchange kinetic energy for magnetic energy, which is what causes the geomagnetic field to deform. The magnetopause represents the boundary where the magnetic forces overcome the dynamic forces of the solar wind. This is a relatively stable boundary requiring an increase in solar kinetic energies, such as occurs with solar flares, to change significantly. The magnetopause is closest at

MAGNETO- TAIL

Figure 4-9. Solar wind. The earth's magnetic field is distorted with its inter- action with the solar wind. Spacecraft Environment 95 about 10 Re, but the tail formed “behind” the earth extends far beyond the orbital distance of the moon (rmmn= 60 Re) Within the magnetopause rests the magnetosphere which defines the cavity containing the (deformed) geomagnetic field.

Geomagnetic Storms. The overall kinetic energy of the solar wind is increased significantly during times of increased solar activity (increases in sunspots and solar flares). When this additional ejected plasma (at speeds up to lo6 km/sec) encounters the earth, it causes a fluctuation of the geomagnetic field as illustrated in Figure 4-10. From several minutes to several hours after a solar disturbance, the addi- tional high-energy particles arrive, causing a compression of the earth’s mag- netic field measured as a sharp increase in the field strength at the earth’s sur- face. This sudden commencement phase of the geomagnetic storm is followed by a two- to eight-hour period during which the magnitude of the field strength remains high. After this initial stage, when the solar wind returns to its prestorm intensity, the geomagnetic field drops to an intensity level significantly lower than normal due to currents set up by increased motions of charged particles trapped within the field, as described in the next section. This main phase lasts about 12 to 24 hours after which a recovery phase begins, lasting typically a few days, during which the field gradually dissipates the excess charges and returns to its normal level of intensity. The drastic alteration of the earth’s magnetic field associated with geo- magnetic storms can affect the operation of many space and terrestrial

INITIAL PHASE MAIN RECOVERY +50 y I PHASE u- -- TIME

SUDDEN COMMENCEMENT

-200y -

Figure 4-10. Geomagnetic “storm.” The intensity of the earth’s magnetic field may be greatly affected by major solar disturbances. 96 Introduction to Space Sciences and Spacecraft Applications systems. The ionosphere may also be altered, which can cause intermp- tion of normal communication channels. The increased intensity of charged particles can affect unprotected systems in space and on the ground, causing overloads, processing or other electronic errors, and increased radiation hazards to humans in space. An associated increase in the motions of trapped charged particles, described next, can also cause an increase in atmospheric density which may affect satellite lifetimes.

The presence of areas of radiation around the earth was discovered by Dr. James Van Allen using instruments aboard the early U.S. V-2 test flights and, later, Explorer series satellites. It was found that some of the charged particles present in space were able to enter the magnetosphere and become trapped within the geomagnetic field. The gainfloss mechanics of this phe- nomenon is still not completely understood, but the levels, dispersion, and behavior of the trapped particles have become well described. Early representations of “belts” of radiation surrounding the earth are better described by distributions of particles within the magnetosphere which follow closely the geomagnetic lines of flux. Areas of different concentrations exist due to the types of particles (protons and electrons) and associated energy levels. More energetic particles are generally trapped closer to the earth, and the major concentrations of particles occur around the equator where the minimum value of magnetic field flux occurs. Figure 4-1 1 shows this distribution of trapped particles around the earth in terms of particle types and energy levels (1E2 indicates 1 x lo2 MeV or mega-electron volts of energy). In some regions, the level of radi- ation due to these trapped particles is sufficient to be disruptive to space- craft operations and hazardous to humans. Although caught in the earth’s magnetic field, these particles are far from stationary. Motion of these energetic charged particles is a combina- tion of spiraling around the magnetic field lines while bouncing from pole to pole along the lines. A slower drift around the equator also occurs. When these particles interact with the ionosphere as they spiral in towards the earth in the vicinity of the poles, they produce the auroral phenomena known as the “northern lights” in the northern hemisphere. A more seri- ous effect of these motions is the contribution to atmospheric heating by particle precipitation in the auroral zones and by the electric currents flowing at high and low latitudes. Spacecraft Environment 97

-i- 12

Figure 4-1 1. Radiation belts. Energetic particles are trapped within the earth’s magnetic field.

South Atlantic Anomaly. The off-set geometry of the geomagnetic field, shown in Figure 4-8,produces an asymmetry in the radiation belts which results in an area where the belts are closer to the earth’s surface. This area occurs just off the coast of Brazil in the Atlantic Ocean and is known as the South Atlantic Anomaly. Spacecraft passing over this area experience a significantly increased radiation dosage sufficient to cause concern when operating in this region.

SPACECRAFT EFFECTS The space systems engineer must not only be familiar with the envi- ronment in which a system is to operate, but must also determine how that environment is going to affect the spacecraft physically and operational- ly. The following sections introduce some of the major concerns. In most cases, a spacecraft must have systems or operational methods to deal with these effects.

Thermal Control Spacecraft in orbit receive thermal inputs from many sources including direct exposure to solar radiation, infrared emissions from the earth, and 98 Introduction to Space Sciences and Spacecraft Applications heat generated by onboard systems. Many of these sources may cycle on and off, such as the solar inputs when orbiting from dark to light or from spacecraft systems as they turn on and off. Spacecraft also continuously emit heat, especially toward the “cold” (absorptive) blackness of space. For some applications, certain spacecraft parts may always point toward the sun while others may always point toward space. Within certain defined limits, the craft must be able to balance the temperatures of all its components. Several spacecraft have been lost or impaired due to the freezing of propellant lines or the heating of an electronic component to the point of failure.

Spacecraft Charging

As we have seen, the space environment is full of charged materials, trapped solar particles, and ionized atmospheric constituents as the major examples. Also, as they travel through space, orbiting spacecraft cut through magnetic field lines of flux which tends to make charges want to move. Operation within such an electrically rich and magnetically dynam- ic environment can result in a buildup of charges on the surfaces of the spacecraft to the point at which a current flow or discharge from one com- ponent to another may result in damage to the system.

Radiation

Some of the higher-energy particles present in space can penetrate the skin and components of a spacecraft. These may deposit an electrical charge inside electrical components or may even change the physical structure of materials. Sudden depositions of electrical charge in comput- er logic circuits or memories can corrupt data or even disrupt operation of the satellite if a “phantom command” is sensed from such an event. Buildup of radiation damage to semi-conductor materials can cause com- puter chips to degrade to the point where commands may no longer be able to be processed. Satellite operators have to be especially vigilant for signs of these effects during periods of high solar activity.

VacuumKorrosion

As we have seen, the near-earth space environment is not a total vacu- um. A spacecraft in orbit is constantly being bombarded by atmospheric Spacecraft Environment 99

particles moving at relatively fast velocities. Some of these particles are the monatomic oxygen atoms created by photodissociation. Being a high- ly reactive element, when these atoms strike a spacecraft they tend to combine with the spacecraft materials, creating a potentially damaging corrosion over long periods of time in space. The major problems associated with operating in the near-vacuum environment of orbits around the earth concern the materials used. Some materials may change their physical properties over time through out- gassing-the release of trapped gases within the materials-or from the radiation and/or corrosion damage discussed earlier. Materials in contact in a vacuum can actually transfer atoms across their surfaces and eventu- ally “cold weld” to each other. This may cause severe problems in space- craft with moving parts. Special substances must be used to separate and lubricate moving parts in space, because lubricants commonly used on earth would simply boil away in a vacuum.

Micrometeorites

The earth is constantly being bombarded by small particles from space. Meteor showers occur when these particles encounter the atmosphere and burn up with the friction of reentry. Though extremely small, these parti- cles may have enormous velocity differences with an orbiting spacecraft. Collision with these micrometeorites can pit the surface of a spacecraft and degrade the performance of sensors or systems (such as solar power cells and thermal coatings) and perhaps cause more serious damage.

Man-made Debris

In the short time that man has been able to place objects in space, the debris created by launch vehicles and spacecraft operations has already created a situation of some concern. The U. S. Space Command has iden- tified more than 7,000 artificial space objects of 10 cm diameter or larger orbiting the earth. Almost 6,000 of these are in low-earth orbit, and about half of the count are spent rocket stages and active and inactive satellites. Computer simulations predict approximately 17,500 objects between 1 and 10 cm in diameter along with millions of smaller particles of space debris. The total mass of man-made debris is about 15,000 times the naturally occurring mass concentration around the earth. The relative velocity between objects that may collide in low-earth orbits is around 10 km/sec 100 Introduction to Space Sciences and Spacecraft Applications

(22,000 mph). Such hypervelocity collision with particles as small as 1 cm in diameter can cause significant impact damage, and shielding for parti- cles much larger than this becomes impractical. This subject has become one of great concern in the design of the international space station. It is estimated that at least one impact with an object 1 cm in diameter or larg- er will occw every 2.5 years for a space station sized vehicle, and about 50,000 impacts with particles 0.01 cm or less would occur each year.

As mentioned earlier, the earth’s atmosphere extends thousands of kilo- meters above the surface, albeit in thinner and thinner proportions. Besides the potentially damaging interactions with atmospheric elements, the drag on spacecraft in this region is also not negligible and may affect spacecraft lifetimes and attitude control. This was graphically exemplified by the reentry of the huge Skylab space station after only six years in orbit. Solar effects on the atmosphere (such as from solar flares and changes in the amount of UV radiations) can cause changes in the atmospheric density, affecting the drag on orbiting spacecraft dramatically and unpredictably.

Sensors

Sensors used in spacecraft are usually very sensitive to particular radi- ations. Direct exposure to solar radiation may result in unwanted signal reception and possible sensor damage. If the purpose of a system is to pick up the weak radiations of a faraway planet or star, and the sensor happens to inadvertently “look” in the direction of the sun (or a reflection of some of its energy), the desired signals may be obscured in the sun’s radiations and the sensor itself may be damaged. Additionally, the sensors must be protected from the general effects of the space environment described above or else performance may be adversely affected.

Spacecraft Performance

The operability and performance of a spacecraft may also be affected by the space environment through which it must operate. Remote sensors must peer through an atmosphere of changing properties (heat, cold, rain) and conditions (night, day), and the transmission of information to ground Spacecraft Environment 101 stations is affected by the atmosphere and ionosphere which also fluctu- ate constantly.

Man in Space

All of the above situations must be considered when man is placed into the space environment. In most cases, a suitable environment can be pro- vided within a spacecraft or with the use of a space suit. However, humans are particularly susceptible to high-energy particle radiations in the form of cosmic rays (from distant stars and galaxies), the particles trapped in the geomagnetic field, and those produced by solar flares. These particles have the ability to pass through shielding and can produce cell damage to astronauts, so exposure must be measured and limited. Fortunately, the areas associated with commonly used nonpolar low-earth orbits have low enough radiation levels to be relatively safe with simply a small amount of protection.

REFERENCEWADDITIONAL READING

National Aeronautics and Space Administration, Space Physics Strate- gy- Implementation Study, 2nd ed. Washington: NASA, 1991. Interagency Group (Space), National Security Council, Report on Orbital Debris. Washington: National Security Council, 1989. T. Wilkerson, M. Lauriente, and G. Sharp (Eds.), Space Shuttle Environ- ment. New York: Engineering Foundation, 1985. C. Cochran, D. Gorman, and J. Dumoulin (Eds.), Space Handbook. Alabama: Air University Press, 1985. Tascione, T., Introduction to the Space Environment. Colorado Springs: U.S. Air Force Academy, 1984.

1. Draw a graph showing the estimated temperature (T) and density (p) variations of the sun from 0 to 2 solar radii (R,) based on the informa- tion given in the chapter. Use logarithmic scales for T and p and a lin- ear scale for R,. 2. Starting with a value for the solar constant in the vicinity of the earth of 1,390 W/m2: 102 Introduction to Space Sciences and Spacecraft Applications

a. Compute the amount of power that must be radiated from the sur- face (photosphere) of the sun (E, W/m2) to produce this value at the earth. b. Compute the total power output of the sun (W). c. Using the figures given in the reading and Einstein’s energy/mass relationship, determine the amount of hydrogen being converted to helium to produce this power (kghec). d. Compute how long it will take until the sun uses up half of the remaining hydrogen. e. Using the Stefan-Boltzmann relationship, determine the correspond- ing surface temperature of the sun to produce the power generated. f. Using Wien’s displacement law, compute the maximum radiated wavelength associated with this temperature. 3. If the earth were a blackbody with an average temperature of 300 OK, determine the wavelength of maximum radiation emitted (pm), the power emitted by the surface of the earth (W/m2), and how much of this power would impinge on a spacecraft orbiting at geostationary altitude (W/m2). (rgeOst,= 42,164 km.) 4. Some objects in space are huge emitters of ultraviolet radiation. If a UV frequency of 3 x 10l6 Hz were detected from a suspected black- body, determine the temperature and the amount of power emitted by the surface of this body. PART 2 Spacecraft Applications CHAPTER 5 Communications

One of the first practical uses for systems permanently stationed in space was proposed by science fiction writer Arthur C. Clarke (most commonly known for writing the story behind 2001: A Space Odyssey) over ten years before the first crude satellites achieved orbit. Having a knowledge of orbital mechanics, Clarke realized that a satellite in a geostationary orbit could connect, in many ways, any two points that might fall within the large field of view. In the early days of telephone, it might not have seemed too unusual to imagine a switchboard in the sky allowing easy, instantaneous communications between such locations. Though devoid of operators, hun- dreds of such switchboards exist today and connect millions of places around the globe from the geostationary band above the earth's equator. COMMUNICATIONS THEORY

Figure 5-1 reviews some of the terms that describe the characteristics of electromagnetic waves as discussed in Chapter 4.

wavelength

amplitude

phase

A = wavelength (distance [m,pmn m = meters p = microns = 1O*m

f = frequency (cycles per sec [Hz]) = CIA c = 3 x 108m/s Figure 5-1. Electromagnetic wave. The characteristics of an electromagnetic wave are used to carry information for communications.

104 Communications 105

This chapter concerns electromagnetic amplitude and phase as well as frequencies and wavelengths. Amplitude represents the amount of energy or power the wave contains, and phase indicates where along the wave form we are at any particular time (in degrees or radians as shown). As we will see in this section, electromagnetic waves are used for com- munication purposes to carry information between locations. Before describing this process, however, some additional information on electro- magnetic propagation must be presented. These properties and phenome- na also apply to remote sensing and satellite navigation, discussed later.

Radio Wave Propagation

The electromagnetic spectrum was described in the last chapter in terms of different bands of frequencies with certain properties. The portions of the spectrum we will be concerned with here are the radio and microwave bands which are most commonly used for communication purposes. Fig- ure 5-2 shows the frequencies associated with these two bands and also shows how these frequencies have been further subdivided into bands z RADIO BANDS

wavelength 1M)o m LOO m 10 m 1 m 10 cm I cm 1 mm 1 CLm

MICROWAVE BANDS

wavelength 1 10 cm lcm lmm

Figure 5-2. Communication bands. Certain frequencies have been designated for particular communications uses. 106 Introduction to Space Sciences and Spacecrafi Applications given names like VHF and UHF. Through international agreements, cer- tain frequencies have been assigned to particular uses, such as AM broad- cast and shortwave. These uses have come about mainly due to the prop- agation properties of the frequencies and their usefulness in carrying certain types of information. These points will be further explained in the following sections. Figure 5-3 shows how electromagnetic waves of certain frequencies propagate differently. These propagation characteristics and the reasons behind each are explained as follows:

Direct Waves. Essentially any frequency can be transmitted between two stations with a direct, unimpeded path between. This is known as line of sight (LOS) communications and is used in many practical applications using frequencies in the radio, microwave, and even visible and other electromagnetic bands.

Ground Waves. At frequencies less than a few megahertz, electromagnet- ic energy can interact with the material in the earth and tend to follow the contour of the earth’s surface. This property allows stations to communi- cate over the horizon (OTH) in longer ranges than direct path stations. As we will see, however, these lower frequencies are limited in how much information they can transmit. Frequencies associated with ground waves are commonly used for public amplitude modulated (AM) broadcasts. Fre- quency modulated (FM) radio stations broadcast on higher frequencies that require direct paths between the transmitter and receiver. This is the reason why AM stations can be picked up generally farther away than FM sta- tions. (Note: Modulation techniques are presented later in this chapter.)

Sky Waves. Frequencies higher than AM frequencies lose the ability to follow the surface contour but, up to a certain point, can have an interac- tion with the ionosphere which allows longer range propagation.

Ionospheric Interaction. The ionosphere was described in Chapter 4 as a region of electrically-charged (ionized) atmospheric constituents. Electro- magnetic waves, upon encountering these electrical charges, tend to change their direction of propagation through refraction. If enough charges are encountered, the wave can be turned back downward toward the surface, appearing to be “reflected by the ionosphere at a certain alti- tude. Depending on the wave frequency and the electron concentration of Communications 107 the ionosphere at the time, some waves will propagate farther into the ionosphere before being turned, appearing to be reflected at a higher alti- tude. Knowing the ionospheric makeup at any time, one can determine the frequencies for which refraction of a signal will occur. The critical fre- quency for this type of interaction is given by equation 5- 1.

where Ne represents the electron density (electrons/cm3) at a particular region of the ionosphere. Equation 5-1 reveals the highest frequency that will be reflected (with noml incidence) by a particular region of the ionosphere. Lower frequencies will also be reflected, but higher frequen- cies will not be refracted sufficiently to be turned completely back down toward the surface. As an example, from Figure 4-5 in Chapter 4 we can see that the high- est electron density for the lowest region of the ionosphere (the “D’ region) is about 103 electrons/cm3. From equation 5-1, the highest frequency that this layer will reflect is about 280 kHz. We can determine the highest fre- quency that will be reflected by the ionosphere as a whole by noticing that the highest electron concentration indicated in Figure 4-5, associated with the F2 region, is about 2 x lo6 electrons/cm3. Equation 5-1 indicates that frequencies up to about 13 MHz will be reflected somewhere in the ionos- phere, and frequencies higher than this may pass through these layers and into space. Note that these results were achieved for the average daytime electron concentrations as given in Figure 4-5. Actual concentrations may be slightly different and, of course, conditions change significantly at nighttime. Both these situations affect the critical frequencies and show why continual monitoring of the ionosphere is important. As noted above, equation 5-1 only addresses a signal propagating nor- mal to the ionospheric layers. A signal that encounters the ionosphere at some angle 9, as depicted in Figure 5-3,will encounter more charged par- ticles as it passes through the regions and will have a better chance to be refracted. As a result, higher frequencies than those obtained for normal incidence will be turned depending on this angle, as can be determined from equation 5-2. 108 Introduction to Space Sciences and Spacecrafi Applications

Figure 5-3. Electromagnetic wave propagation. Certain frequencies of electromagnetic energy propagate differently.

The ranges achieved depend on a number of factors including frequen- cy, angle of incidence, and transmitted power (to be discussed shortly), as well as time of day (or night) and other factors which may affect the prop- agation of the signal. A powerful enough signal may, after refraction by the ionosphere, reflect off the earth’s surface back toward the ionosphere to be refracted again. These multiple-hop transmissions can achieve tremendous ranges and allow essentially worldwide communications. Fre- quencies used for this type of communication make up the shortwave band which has been used for long-range communications since Marconi explored the possibility of such use at the beginning of this century.

Space Waves. Above the critical frequency (generally above 30 MHz), electromagnetic waves are not affected by the ionosphere enough to be deflected back down to the surface. Communications using these fre- quencies are all line-of-sight (LOS) communications because the trans- mitting station and receiver must have a direct path between the two.

Antenna Theory

Electromagnetic waves that propagate through space originate from, and are received by, some sort of antenna structure. Power is delivered to an antenna at the desired frequency, causing electron movement in the structure. The movement of electrons causes an associated electromag- Communications 109 netic wave to be formed which propagates from the antenna in a certain pattern depending on the physical characteristics of the antenna.

Antenna Patterns. An antenna pattern is a geometrical plot that repre- sents the way in which an antenna propagates electromagnetic energy. In most cases, the pattern is the locus of points that represent the range at which the transmitted power is decreased decreased equivalently by 3 decibels (dB) of its original value (or some normalized equivalent value). The relationships between decimal values and decibel values are given by:

and

-x dB x=lO 10 (5 - 4) The second of these equations shows that -3 d3equals one-half in dec- imal value. This indicates that the locus of -3 dB points of the antenna pattern represent the half-power locus delineating where the signal has decreased to one half of its originally transmitted power level.

Isotropic Radiator. An isotropic radiator is a (theoretical) structure from which electromagnetic radiations would propagate in all directions at equal power strengths. The antenna pattern for such a radiator would look like a perfect sphere of half-power points emanating uniformly in all directions from a central point.

Anisotropic Radiators. In many cases, it is desired to concentrate trans- mitted radiations in a particular direction or pattern in order to reach cer- tain areas and/or certain power levels in those directions. All practical antennas are designed, depending on their specific purpose, to transmit (and receive) signals in such a way. For instance, the radio antenna on your car must be able to receive AM and €34 signals from nearby radio stations in any horizontal direction (since you may be making turns as you drive), but because no stations currently exist in space or underground, the antenna does not need to pick up signals from these directions. The famil- iar pole antenna on your car has an antenna pattern that looks like a doughnut (when looking from above or below), indicating that the anten- 110 Introduction to Space Sciences and Spacecraft Applications na concentrates its attention horizontally with decreased receiving capa- bilities in other directions.

Parabolic Dish. In many applications, it is desired to concentrate a signal as much as possible in a single direction. For these applications, a para- bolic dish antenna is most commonly used. Figure 5-4 depicts a parabol- ic dish with a representative antenna pattern, and also delineates some of the terms with which we will be dealing. When transmitting, the signal is either emitted by, or reflected off, a structure at the focal point of the antenna paraboloid. The parabolic shape of the antenna reflects the signal mainly along the boresight, concentrating most of the signal power in this direction. When receiving, the incoming electromagnetic wave “reflects” off the dish and is concentrated onto the focal point of the parabola where (either by pickup or secondary reflection) the signal is delivered to the receiving components. An antenna has the same propensity for receiving a signal from a certain direction as it has for transmitting in that direction, as quantified in the following section.

Antenna Gain. The ability for an antenna to concentrate a signal in a cer- tain direction is given by the antenna’s gain in that direction. This prop- erty is dependent on the size and shape of the antenna structure and on the frequency of radiation as well. We will be dealing mainly with parabolic dish antennas in this chapter, and a simple relationship for gain in the boresight direction for such an antenna is given by equation 5-5.

PARABOLIC DISH REFLECTOR

I RECEIVEW BORESIGHT TRANSMllTER VLP 2

ANTENNA PAlTERN (locus of 1/2 power)

Figure 5-4. Parabolic antenna. Parabolic antennas are used to concentrate transmitting or receiving energy in a particular direction. Communications 111

where h is the wavelength of the signal and Ae represents the efsective antenna area which, for a parabolic dish, is given by equation 5-6.

where D is the dish diameter (as depicted in Fig. 5-4)and p is an antenna eficiency factor (typically around 0.5-0.8). Notice that using the same units for A, and h results in a dimensionless value for the gain. As you can imagine, directions away from the boresight of a parabolic dish antenna reduce the effective area that can be seen, decreasing the associated gain of the antenna in that direction. The antenna pattern shown for the parabolic dish in the sketch also represents the antenna’s gain. Though it is easy to imagine this as similar to a flashlight beam emanat- ing from the antenna (which it is), it must be realized that this same gain applies to the antenna when it is receiving signals as well as when trans- mitting them. Parabolic antennas send and receive signals best from the boresight direction and do not receive signals as well from other direc- tions. Numerically, the same relationship that gives the boresight gain for a transmitting dish (GT), also applies to the antenna’s ability to receive signals from the boresight direction, called the receiving gain (GR).

Power Budget

The amount of power that arrives at an antenna, and the ability of the antenna and receiver to pick up and recognize that signal, determines if successful communications will occur. This is determined by a link (or power) budget analysis of the specific system, as depicted in Figure 5-5. A certain amount of transmitter power (PT,usually described in watts [w] or decibel-watts [dBW]) is delivered to the antenna. If this were an isotrop ic radiator, the power would be radiated uniformly in all directions and at some distance (R)from the antenna the amount of power detected would be: 112 Introduction to Space Sciences and Spacecrajl Applications

Figure 5-5. Link budget. A certain minimum amount of power must be received for successful communications. because the power would be spread out throughout the 47c radians of the sphere around the isotropic radiator. Notice that this is an inverse square relationship which describes the loss of power of any electromagnetic sig- nal propagating through space. The gain of the antenna concentrates the power in the boresight direc- tion and represents a multiplicative factor for the power in that direction. At a distance of R along the boresight direction from a parabolic antenna, the power would be:

The above relationship gives the power per unit area at range R. The amount of that power that falls on the receiving antenna depends on that antenna's effective antenna area A,, the expression for which was given in equation 5-6.

Relating this result to the form of equation 5-5 for the receiving anten- na gain (GR), we can write: Communications 113

We can define a term known as thefree space loss (Ls) which repre- sents the losses due to the spreading out of the signal through space between the transmitter and receiver:

Ls=(h)4xR (5 - 7)

Incorporating the free space loss term, our equation simplifies to:

Finally, this expression usually incorporates a term that takes into account a multitude of other signal power losses due to factors such as the atmosphere, ionosphere, electromagnetic interference, and the like. This term is called the incidental loss term (L,) which, as a loss term, appears in the denominator of the above equation. The final form of this equation is called the link (or power) budget which gives a value for the power received by a station at a distance R from a transmitter:

In many cases this calculation is made using decibel values which would, for the above relationship, be written as:

where each term is either specified in or converted to dB using equation 5-3.

Noise. The amount of power received may still not determine if the receiv- er will understand the transmitted information, which is required for suc- cessful communications. The reason for this is the presence of noise which can interfere with the transmitted signal. There are many sources for this noise including the transmitting and receiving equipment as well as natur- al causes such as the sun, earth, atmosphere, and even other celestial bod- ies. The amount of noise a source produces is represented by the source’s equivalent noise temperature (Teq).For our purposes, we will assume that 114 Introduction to Space Sciences and Spacecrajl Applications the receiving equipment is the major source of noise and that its noise tem- perature will be known or given. The resulting noise is given by:

(5-10) where k = Boltzmann's constant (k = 1.38 x W/"K Hz) and bw rep- resents the bandwidth or the range of frequencies the receiver has been designed to pick up. Less noise may be experienced by using a small oper- ating bandwidth, but in all practical cases the desired information signal is spread out over a range of frequencies dictating what the minimum bandwidth for a receiver must be.

Signal-to-Noise Ratio. Successful communication systems (known as links) are designed to operate under all expected conditions of losses and noise. The measure of this is given by a receiver's signal-to-noise ratio, which is given by equation 5- 11:

SIN = -PR (5-11) NO A receiver will have a minimum signal-to-noise ratio specified for acceptable operation. If the received power decreases or noise increases such that the actual signal-to-noise ratio is less than the minimum speci- fied for a receiver, the information in the carrier signal may not be intelli- gible or the carrier signal may not be detected at all.

Example Problem:

A geostationary communications satellite has a traveling-wave- tube (TWT) amplifier with a maximum power output (PT)of 9 W and transmits via a 5 m diameter parabolic dish antenna with an effi- ciency of 0.6. The antenna is boresighted on a ground station using a 30 m diameter parabolic dish antenna (p = 0.5). The receiver band- width is 1 MHz wide around the 6 GHz down-link frequency, and the receiving components are cooled to maintain an equivalent tem- perature of 300 OK. Slant range (distance between antennae) is 40,000 km and the incidental down-link power losses (L,) are 1 dB. Determine the signal-to-noise ratio of the receiver. Communications 115

Solution:

Ae(,, = 11.78 m2 = 10.71 dB Ae(,, = 353.43 m2 = 25.48 dB GT(*)= 59,217.63 = 47.73 dB GR(')= 1,776,528.8 = 62.5 dB LS(*)= 1.011 x IO2' = 200.05 dB PR(3)= 7.44 x W = -81.3 dBW No = 4.14 x W = -143.83 dBW PR/N,= 1.8 x IO6 = 62.55 dB - -. _. ___ - _.. - __ Noteq (I) h = c/f = 0.05 m (2) R and h must use same units (3) I dB = 1.26 decimal

MODULATION

Modulation is the means by which information is impressed on elec- tromagnetic radiation so that it may be transmitted from one location to another. There are at least two different components to a communications signal: the carrier wave and the baseband signal. The baseband signal represents the information desired to be transmit- ted. In the case of a radio station, this may be the audio frequencies (of a few thousand hertz) of the DJ's voice or the music of a record. These are good examples of analog signals which are the raw, continuous wave-like frequencies of sound or light. Television stations transmit higher baseband frequencies that are used by a TV set to illuminate the tube in a certain way to produce a picture. Banks and businesses (and satellites) may con- vert their baseband information (numerical transactions, written text and pictures, remotely sensed frequencies) into digital signals (strings of ones and zeros) which then represent the original information which may allow improved transmission of the data. Currier waves, as the name implies, are electromagnetic signals used to carry the baseband information between the transmitter and receiver. Car- rier waves are used because the baseband signal may not have sufficient propagation characteristics for successful communications. For instance, audio frequencies are attenuated severely by the atmosphere, limiting their useful ranges (can you imagine radio stations that simply blasted their music from loudspeakers all over the city?). The particular carrier wave frequency may be chosen primarily for its propagation characteris- 116 Introduction to Space Sciences and Spacecra# Applications tics. For example, if communications with a satellite in orbit is desired, the signal frequency must be above the critical frequency of the ionos- phere or the signal will not pass through into space. As important as its propagation characteristics, carrier waves are also chosen according to the amount of baseband information that must be trans- mitted. Only so much information can be impressed upon a carrier wave and still be recovered when the signal is received (and demodulated) by the receiver. As a rough rule of thumb, it can be considered that a carrier wave can only be impressed with basebandfiequencies up to about 10% of the carrier wavefrequency. So, to transmit the human voice (around 4 kHz) a minimum carrier frequency of about 40 kHz would be required. As mentioned above, the baseband information must be impressed onto the carrier wave for propagation between stations. In all cases, some char- acteristic of the carrier wave, such as amplitude or frequency is modified to represent the baseband information whether it is in analog or digital form. The following sections describe the general modulation techniques using analog baseband signals. Digital communication techniques are very similar to these analog techniques and will be presented later in this chapter after digital signals themselves are more fully described.

Amplitude Modulation (AM)

In amplitude modulation schemes, the baseband signal is used to change the amplitude of the carrier wave in such a way as to represent the baseband information. Figure 5-6 shows how an analog baseband signal modifies a carrier wave to produce an AM signal. Notice that the frequency of the carrier wave has not changed. Only the amplitude varies and it does so in exactly the same manner as the base- band signal. The AM signal will propagate between stations with the char- acteristics of the carrier wave, but if we only look at its amplitude, we can see the characteristics of the information impressed upon the signal.

Frequency Modulation (FM)

In frequency modulation, the frequency of the carrier signal is modified by the baseband signal, as illustrated in Figure 5-7. Note that the ampli- tude of the signal remains constant. The frequency variation from the known carrier frequency is directly proportional to the frequencies of the baseband (information) signal. Communications 117

BASERAND (INFORMATION) SIGNAL

CARRIER WAVE

AMPLITUDE MODIILATED SIGNAL

Figure 5-6. Amplitude modulation. The amplitude of the carrier wave is changed in a way to represent the information to be transmitted.

CARRIER WAVE

FREQUENCY MODIIIATED SIGNAL

Figure 5-7. Frequency modulation. The frequency of the carrier wave is changed in a way to represent the information to be transmitted.

Phase Modulation (PM)

Information can also be impressed upon a carrier signal by using the baseband signal to modify the phase of the carrier wave, called phase modulation. The rate at which the carrier phase is changed, and/or the location along the wave at which the change occurs, can be used to repre- sent the baseband signal. An example of a PM signal is shown in Figure 5-8. Notice, in this case, that both the amplitude and frequency of the signal remain constant and that only the phase changes, as indicated by the arrows in the figure. Phase 118 Introduction to Space Sciences and Spacecraft Applications

CARRIER WAVE

PHASE MODUMTED SIGNAL

Figure 5-8. Phase modulation. The phase of the carrier wave is changed in a way to represent the information to be transmitted. modulation is not used much in conventional (analog) communications systems, but is widely used in digital communications schemes employed by many satellite systems.

Demodulation

Demodulation is defined as the recovering of the baseband signal from the received modulated signal. The basic process used is simply the opposite of that used to modulate the carrier signal in the first place. For AM signals, a device called an envelope detector senses the change in amplitude of the received signal and reconstructs the baseband signal from this. Similar meth- ods are used by FM demodulators (frequency detectors) and PM systems (phase detectors). It is very important that the receiver be able to reproduce the carrier wave exactly, as this is used by the demodulator to “subtract” this component from the received signal (when you “tune” in a station you are adjusting the receiver’s oscillator to match the desired carrier frequency). Though demodulation may sound simple, the procedure is complicated by the fact that the received signal is usually very weak, with the infor- mation-carrying portion of the signal (PR)possibly weaker than the noise signal (No). Also, due to its transmission through the atmosphere (and per- haps the ionosphere), the signal may be significantly corrupted (changed by interaction with the elements) which may affect proper recovery of the information signal.

DIGITAL COMMUNICATIONS

Digital communications offers some benefits over analog communica- tions, mainly in terms of the transmission, reception, and handling tech- niques of the digitized information. Since more and more systems are Communications 119 incorporating solid-state technology (which is itself a digital medium), the use of digital signal processing has increased as well. Some familiar examples are CDs (compact disks), on which music is stored digitally and reconverted into analog sounds by your player, and telephone links (fiber- optic digital telecommunications). Even radio and television stations are beginning to digitize their signals. In this section, we will discuss how a baseband signal (information) is digitized, what the characteristics of such signals are, and how the resulting signal is modulated and demodulated.

Basic Binary

Digital signals are merely representations of the original information using a simple (binary) coding system. The invention of the transistor in the 1950s began the interest in this field. With the transistor, it was possi- ble to represent a situation (“true” or “false,” for instance) by indicating either “true” by allowing current flow through the device (device “on”), or “false” with no current flow (device “off’). Since only two situations are possible, a transistor is a binary (2-ary) device, and in binary language, “0n” could represent a digital “1” and “off’ a digital “0.” A single transistor can only represent two situations, or states, at any one time (either 1 or 0 could be indicated). If two transistors are combined simultaneously, four different states could be represented at any one time (0 and 0,O and 1, 1 and 0, or 1 and 1, respectively, on the two transistors). Three transistors allow eight different states to be represented, four tran- sistors 16 states, and so on. Table 5-1 shows the possible states combina- tions of three and four transistors. (Note: Each 1 or 0 that represents a pos- sible state for a transistor is known as a bit.) Notice that, in going down the table, to get from one state to the next on the list you need only add (in binary) a “one” to the preceding value. There is an entire branch of mathematics concerned with the manipulation of bina- ry numbers, but for our purposes we need only the simple relationships which relate the number of states that a certain number of bits can represent:

2” = k (5- 12) and

n = 1.44271n (k) (5-13) 120 Introduction to Space Sciences and Spacecrafi Applications

Table 5-1 Number of Bits vs. Number of States

3 BITS 4 BITS 000 0 000 001 0 00 1 0 10 0 010 0 1 1 0 011 1 00 0 100 1 01 0 10 1 1 10 0 110 1 1 1 0 111 8 STATES 1 000 1 00 1 1 010 1 01 1 1 100 1 10 1 1 110 1 111 16 STATES where n is the number of bits that can represent k different states. For instance, a typical personal computer has a single chip processor with a 16- bit capability allowing 65,536 different states. This means the computer can perform 65,536 different operations, with each letter, number, symbol, or function counting as, at least, one operation. This really is quite a capability! As mentioned above, each 1 or 0 is known as a bit. In computer lingo, four bits put together are called a nybble, and eight bits constitute a byte. A word is a combination of bits (or nybbles or bytes) used to represent something of value to the user. Of course, everyone using the same infor- mation must know how many bits make up each word, as well as how the bits are arranged or ordered, or else the information would just be a mean- ingless stream of ones and zeros.

Analog to Digital (ND) Conversion

The following sections discuss how an analog signal (like music, a voice, or even visible frequencies) is converted into groups of digital Communications 121

111

110 101 100

01 1 01 0 001 OOO

Figure 5-9. A/D conversion. An analog signal may be sampled periodically and given digital values used to represent the original signal. words that can represent the original signal. Figure 5-9, which will be referred to in the following discussions, depicts a single sinusoid of an analog signal being digitized.

Quantization. Quantization is the method of assigning a range of digital values to represent the original signal. Two factors limit this process: the range of signul values that must be covered, such as amplitude range or fre quency range, and the number of bits that the system can use to represent these values. For the signal in Figure 5-9, the system must be able to rep- resent the signal from peak to subsequent peak of the sinusoid shown. With only a single bit we could indicate that the sinusoid was either above (with a “one”) or below (with a “zero”) the zero amplitude line, but nothing more. With two bits we could split the above and below portions of the sinusoid in half and could indicate (still crudely) how far above or below the signal was at that point. The signal shown in the figure has been split into eight quantization levels (known as Q-ZeveZs) giving a better ability to represent the signal amplitude. As you can imagine, with an unlimited number of levels we could represent exactly where above or below the line the signal was at any time. Unfortunately, there are limitations to the num- ber of bits that a system can use (as we will discuss shortly) and we find that we must try to represent the signal with as few bits as possible. We can use equations 5-12 and 5-13, where k now represents the num- ber of Q-levels and n the number of bits, to determine the number of bits required or the number of levels available given one of the parameters as a system requirement. 122 Introduction to Space Sciences and Spacecraft Applications

Coding. The method of assigning a particular bit word to each quantiza- tion level is called coding. (Coding for security purposes is another sepa- rate process, the mechanics of which are obviously more complicated than the simple digitization coding discussed here.) Because the signal in the figure has been split into eight Q-levels, three bits are required to describe each level separately. The coding used is the specific order in which the eight different three-bit words are assigned to each level. For the signal shown in Figure 59,000 has been assigned to the bottom level and a bina- ry “one” is added to get the coding for each subsequent level.

Sampling. Quantization values are assigned to a signal often enough to allow the receiver of the information to reconstruct the original signal (sort of by simply connecting the dots indicated by the digital informa- tion). The frequency at which a signal is assigned quantization levels (sampled) is called the sample period (t,). It should be obvious that if a signal were sampled too infrequently, not enough information would be available to reconstruct the original signal. However, as with quantization, sampling too often may result in too large a number of bits for the system to be practical. It has been shown that if a signal is sampled at least twice during each sinusoid, there will be enough information to reconstruct the signal. In terms of frequencies:

f, = > 2fm,, (5-14) where fs represents the sampling frequency (sampleshec), and fmaxrepre- sents the highest frequency in the range of signals being sampled. Note that the sample period is simply the inverse of the sampling fre- quency as shown in equation 5-15. 1 t =- (5 - 15) fs Transmission Rate. The transmission rate represents the number of bits per unit time that a digital system creates. Since a sample of n bits is being taken every t, seconds, and since f, = 1/ t,, the transmission rate (TR) can be found from:

n TR = - = n x f, (bps) (5 - 16) f, Communications 123 where bps stands for bits per second. The transmission rate is an impor- tant parameter of a digital communication system, as it is a measure of the capabilities of the system to transmit information. Most modern space- craft use digital communication methods, and in many cases the trans- mission rate is a major factor limiting the system capability. Examples are telecommunications satellites which can only handle so many telephone calls at a time or remote sensing systems that can only observe so much information due to data storage and transmitting limitations. The major limiting factor to transmission rate (and subsequently, the limit to the sam- ple rate and number of Q-levels) is the carrier wave’s ability to transport digital information, as described in the next few sections.

Example Problem:

Satellite transmission of telephone calls allots a bandwidth (range of frequencies) of 4 kHz (associated with the minimum bandwidth required to reproduce the human voice) for each call. If eight Q-lev- els are used to digitize each call, determine the minimum sampling rate (frequency), sampling period, and transmission rate associated with one conversation.

Solution:

fsmin = 2 fmax = 8,000 samples/sec t, = l/f, = 0.125 x lod3sec n = 1.4427 In(k) = 3 bits/sample TR = n x f, = n/t, = 24,000 bits/sec (bps) = 24 Kbps

Digital Baseband Signals

The output of an AD converter is a string of “one” and “zero” values representing the bits assigned to each successive sample of the original signal. These values are usually output as voltages, where a positive volt- age may correspond to a “one” and either a negative or zero voltage would represent a “zero” bit. Figure 5-10 shows these two common methods of representing a digital string of values. 124 Introduction to Space Sciences and Spacecraft Applications

0 1 0 0 1 1 0 1 . . DIGITALSTRING

RETURN &&b ’&&b TO(RTz) ZERO

TO ZERO ( NRZ ) -V k=f-=tb 2tb 5th 4tb 5tb 6tb 7tb 8tb NoN-RETURN Figure 5-10. Digital baseband signals. The “ones” and “zeros” of digital baseband signals are usually represented by voltage levels held for a specified period of time.

The time that the specific voltage is held to represent each bit is called the bit period (tb), as indicated in the figure. It is desirable to represent each bit as long as possible, as this would allow a receiver more time to recognize each bit as it is received. But for real-time (continuous) com- munications, the number of bits representing each sample must be trans- mitted before the next sample bit stream comes up. To ensure this, the bit period must be less than or equal to the sample period divided by the num- ber of bits per sample: t tb = 5 2 (sec/bit) (5- 17) n

The pattern of this voltage output now represents the information to be transmitted and will be used as the baseband signal to modulate the carri- er wave in the methods described next.

Digital Modulation Methods

The digital baseband signals that are the product of A/D conversion are used in methods similar to those discussed earlier in analog modulation schemes to impress information onto a carrier wave for transmission between stations. It was mentioned earlier that analog baseband signals were limited in frequency to about 10% of the carrier wave, and we find that digital signals have similar limits on the number of bits that a carrier wave can transport. We can simply consider a digital baseband signal as a continuous waveform with a period equal to two bit periods, as depicted in Figure 5-1 1 which shows the first few bits of a NRZ waveform. Communications 125

CI-) f ‘‘cant" = 1/2t b fb “cont”

The maximum transmission rate a digital system could have, based on a given carrier wave frequency (fc), can be found from the following rela- tionships:

(5- 18) where fb“c,,nt” = % tb, representing the equivalent “continuous’’ frequency of the digital waveform. Once a maximum transmission rate is deter- mined, the bit period, sampling frequency, and sample period can be established.

Example Problem:

A communications satellite uses a 6 GHz down-link frequency to transmit digitized telephone calls. Using the results of the last exam- ple problem, determine the maximum transmission rate and the number of calls that may be transmitted simultaneously.

Solution:

From the last example problem we found that for each call:

fSmin= 8,000 samples/sec t, = 0.125 x lop3sec n = 3 bitdsample TR = 24 Kbps 126 Introduction to Space Sciences and SpacecrafrApplications

The maximum transmission rate for the satellite is:

TR,,, = fJl0 = 6 x lo8 bps = 600 Mbps

which indicates that simultaneously the link could handle:

TR,,/TR (one call) = 25,000 calls

assuming that the baseband information for each call can be combined into a single continuous bit stream. The methods used to accomplish this are beyond the scope of this text, but notice that the maximum peri- od for each bit on the carrier wave must be much less than that for each call separately:

tb (one call) = t$n = 4.2 x sechit tb (link) = 1.4 fb”cont~9= l/TR,,, = 1.7 x sechit This result reveals that less than two nanoseconds can be allotted for each bit transmitted!

The following sections describe the basic digital modulation methods and their similarity to the analog modulation techniques discussed earlier.

Amplitude Shift Keying (ASK). In this modulation method, the carrier wave is either transmitted (representing a digital “one”) or not transmitted (signifying a digital “zero”) each bit period. The relationships between the

CARRIER WAVE

Figure 5-12. Amplitude shift keying modulation. The amplitude of the carrier wave is used to represent the digital baseband signal. Communications 127

baseband signal, carrier wave, and the resulting ASK signal are shown in Figure 5-12.

Frequency Shift Keying (FSK). In this modulation scheme, two carrier frequencies are used, one to represent a transmitted “one” and the other a digital “zero.” An example of an FSK signal is shown in Figure 5-1 3.

0 1 0 0 1 1 0 1 - DIGITALSTRING

FREQUENCY SfflFT KEYED SlCNAL

Figure 5-13. Frequency shift keying modulation. The change in frequency of the carrier wave is used to represent the digital baseband signal.

Phase Shift Keying (PSK). In phase shift keying, a change in a bit state (from “zero” to “one” or from “one” to “zero”) is indicated by a change in the phase of the carrier wave. Notice that if there is no change in the bit pattern, the phase of the carrier signal doesn’t change until there is one, as shown in Figure 5-14.

PHASE SHIFT KEYEU SICNAL

Figure 5-14. Phase shift keying modulation. The change in phase of the carrier wave is used to represent the digital baseband signal.

An important point in PSK systems is ensuring that the receiver knows what the starting bit is supposed to be, or else the reconstructed digital string will be opposite that transmitted. This is usually done by first send- ing a code which, whether received correctly or not, will indicate what the first bit in the message should be. 128 Introduction to Space Sciences and Spacecraft Applications

Demodulation. Demodulation methods are similar to those described earli- er for analog signals in their use of envelope, frequency, and phase detectors. For digital demodulation, the demodulator must only detect the presence of a transmitted “one” or “zero” which is relatively more simple than recover- ing the actual analog baseband signal from the received signal (plus noise). This is one of the advantages of digital communications over analog systems, as it allows a successful link with lower minimum signal-to-noise ratios. An important factor for successful use of digital signals for communi- cations is ensuring that everyone on the link is synchronized in time. Mod- ification of the carrier wave by the transmitter modulator only occurs each bit time (which, as we’ve seen, can be measured in nanoseconds), and dig- ital demodulators are looking to detect changes in the received signal at these times only as well.

COMMUNICATIONS SYSTEMS

Figure 5- 15 displays the components of a typical communication system. In the transmitter, the source represents the originator of the information signal (FLtl).This may be the microphone of a radio or telephone; the video receivers of a television camera; the antenna of a remote sensor; or whatever is picking up or originating the desired information, turning it into electromagnetic signals suitable for transmission. If the signal is to be digitized, an An> converter would be part of the preparatory electronics that make a signal ready for modulation onto the carrier wave which occurs within the modulator section of the transmitter. The anzpZ$er

TRANSMITTER

MODULATOR H AMPLIFIER H ANTENNAF CHANNEL+ x(t)

RECEIVER

Figure 5-1 5. Communications systems. Many of the common components of a typical communications system are depicted. Communications 129 boosts the signal up to the desired transmitter power (PT) and delivers this to the antenna from which the electromagnetic signal (XIt1)will propa- gate. The channel represents the medium through which the signal travels between stations. (For a radio station, this medmm would be the atmos- phere between the station and your car. The channel for telephone or cable TV is the line coming in your house.) On the receiving end, the antenna picks up the signal (with its associat- ed noise) from the channel. Since the power level received is usually quite low compared to that which was transmitted, the signal is delivered direct- ly to an amplifier to boost the signal strength. The front-end electronics attempt to filter out some of the signal noise before delivering the signal to the demoduZator; which recovers the baseband signal from the carrier wave frequency. (Note: As was mentioned earlier, the receiver electronics contribute significantly to the overall equivalent noise temperature of a communications link. To minimize this contribution, many receiving sta- tions cool their front-end electronics, some down to near absolute zero using liquid nitrogen, to improve the received signal-to-noise ratio). Finally, the back-end electronics reproduce the transmitted information in its original form (whether by speaker, TV tube, teletype, etc.). It is here that the digital-to-analog (D/A) converter will reconstruct the original sig- nal if it was first digitized before transmission.

REFERENCES/ADDITIONAL READING

Pritchard, W., Suyderhoud, H., and Nelson, R., Satellite CommunicationSys- tems Engineering, 2nd ed. Englewood Cliffs: Prentice-Hall, Inc., 1993. Feher, K., Digital Communications, Satellite/Earth Station Engineering. Englewood Cliffs: Prentice-Hall, Inc., 1983. Roden, M., Analog and Digital Communication Systems. Englewood Cliffs: Prentice-Hall, Inc., 1979. Gagliardi, R., Introduction to Communications Engineering. New York: John Wiley & Sons, 1978.

1. A ship at sea wishes to communicate with a shore station over the hori- zon using its HF radio. Using the F2 region of the ionosphere for a “reflection” point between the stations, it is determined that the inci- 130 Introduction to Space Sciences and Spacecraft Applications

dence angle must be 30" to get the range required. Determine the high- est frequency that the ship can use to transmit its information. Would this be a sufficient carrier frequency to transmit a 3 MHz baseband sig- nal? (Explain.) 2. A geostationary communications satellite has a traveling-wave-tube (TWT) amplifier with a maximum power output (PT)of 9 W and trans- mits via a 5 m diameter parabolic dish antenna with an efficiency of 0.6. The antenna is boresighted on a ground station with a 30 m diameter par- abolic dish antenna (effkiency = 0.5) with a receiver cooled to maintain an equivalent temperature of 300" K. Slant range (distance between antennae) is 40,000 km, and the down-link frequency is 6 GHz. The inci- dental down-link power losses (LI) are 1 dB. Do this problem twice, once using the decimal form of the equations and values, and then again using dB values to perform the calculations. a. Compute the gains of the transmitting antenna (GT) and the receiving antenna (GR). b. Compute the free-space loss (Ls). c. Determine the power received by the ground station (PR). d. What is the received signal-to-noise ratio (PR/No)? e. Show that the values obtained in part (d) above, using the decimal approach and the dB approach, are equal. 3. Voyager 2, sent out to visit the outer planets, is in a hyperbolic trajec- tory which will carry it out of the solar system. Scientists are interest- ed in monitoring Voyager 2 for as long as possible to see if there is an interface between the space in the vicinity of our sun and true interstel- lar space (the heliopause), and what the interstellar space may be like. Voyager 2 has a 3.7-meter diameter parabolic dish antenna with an effi- ciency factor (p) of 0.61. The onboard 21.3 W transmitter sends data back towards the earth on a down-link frequency of 8.415 GHz (X- band). Signals from Voyager 2 are picked up by NASA's Deep Space Network (DSN) parabolic antennas that are 64-meters in diameter (p = 0.42) with receivers that are cooled down to 28.5" K. Incidental losses are estimated to be 1.549 dB and the minimum signal-to-noise ratio that the DSN receivers can tolerate is 2.05 x lo5. a. Determine the signal power (in watts) received from the spacecraft by the DSN at the time of minimum S/N reception. b. Compute the distance from the earth at which contact will be lost (in terms of A.U.s where 1 A.U. = 149.5 x lo6 km). Communications 131

c. Will we still be in contact with the spacecraft when it flies beyond the farthest reaches of Pluto? (apluto = 5.9 x lo9 km, epluto= 0.25) 4. An FM radio station wishes to transmit baseband frequencies from 20 to 20,000 Hz in order to broadcast good quality music. The station wants to transmit digitally (by first converting the baseband signal to a digital signal) using byte-sized digital words. a. Determine the number of Q-levels the signal may be assigned values from. b. What is the minimum sampling frequency (and associated sampling period) that may be used? c. Determine the bit period, the transmission rate, and the minimum carrier frequency that might be used to transmit the music. d. Can the current frequency range assigned to FM broadcasting be used by this station? (Explain.) 5. The Army and Marine Corps would like to have the capability for satel- lite communications with ground forces using a portable transceiver. Such a radio would have a small (1 meter) diameter parabolic dish antenna with a low efficiency (0.5). Maximum transmitted power would be 2 watts. The soldier would like to be able to transmit to a geostation- ary satellite down to 5 degrees elevation angle (v = 5" [described in Chapter 61, klmt-max= 41,126.7 km). The satellite has a 5-meter diam- eter parabolic dish (efficiency = 0.9) and uses deep space as a heat sink to keep the receiver cooled to 75°K.Up-link frequency is 30 GHz and worst case (\v = 5") incidental losses are 4 dB. Consider the antennae as being able to boresight on each other during communications, and deter- mine the minimum signal-to-noise ratio the satellite will have to be able to handle for a successful link. While showing your calculations for the link budget, put a box around the following parameters: CHAPTER 6 Remote Sensing

On August 7, 1959, sent back the first satellite-produced pic- ture of the earth taken from an altitude of about 20,000 km in space. Although grainy and lacking detail, it gave a spectacular view of swirling cloud formations over the oceans and landmasses of a large portion of the earth. Perceiving the immense value of observations from space, planners quickly began to take advantage of the benefits of this perspective view, and a few months later, on April 1, 1960, TIROS I was launched. TIROS I ushered in an age of continuous operational meteorological satellites (often referred to as weather satellites) by returning thousands of pictures of the earth during its two-and-a-half-month lifetime. These and many other pio- neering satellites demonstrated the basic advantages of satellite remote sensing: the relatively large (compared to airborne perspectives) amount of the earth in view at any time; the rate of coverage possible from a space- based platform; and the ability to view even the remotest locations on the earth’s surface. For example, a typical polar orbiting weather satellite at an altitude of 830 km circles the earth every 102 minutes, moving with an earth velocity of nearly 6,600 dsec, viewing thousands of square kilome- ters each second as its sensors scan cross-track from horizon to horizon. Two of these operational meteorological satellites obtain a global view of the earth twice a day, providing government and private weather services with surface and atmospheric data that improve and extend weather pre- diction. A stark contrast and testimony to the value of satellite weather observation is the accuracy and timeliness of observations of severe storms, such as hurricanes, compared to presatellite times when we relied on data from ships of opportunity or sparse information from weather air- craft designated to search for hurricanes. Although property damage still occurs, the advanced warning provided by timely space-based weather observation mitigates damage and has sharply reduced loss of life. There are a multitude of remote sensing satellites that enjoy the cover- age benefits offered by the vantage point of space. In addition to environ-

132 Remote Sensing 133 mental monitoring satellites, there are a host of fine spatial resolution satellites that owe their heritage to the pioneering work done by NASA’s Landsat and Seasat satellites. These satellites employ electronic cameras and special radars to observe the earth and its oceans in great detail for military, scientific, and commercial purposes. Remote sensing from space can be categorized in a variety of different ways, such as the type of sensor employed, the application purpose, and/or the sponsorship of the mission. These themes will be expanded in the sections that follow but are first outlined in this introduction.

Sensors

Sensors can first be categorized as being passive or active. A passive sensor receives information from the earth and the atmosphere. A camera is a good example of a passive sensor relying on solar illumination of the object(s) in view to provide input to the camera. By contrast, an active sensor provides the source of energy used to view the area of interest. Radars and lasers are examples of active sensors. Both passive and active sensors operate over a wide portion of the elec- tromagnetic spectrum, The precise frequencies or wavelengths chosen are a function of first, the application, and second, the influence of the propaga- tion path between the satellite and the earth. The signatures provided by the earth and its atmosphere result from highly complex interactions of induced, in the case of an active sensor, or natural, in the case of a passive sensor, electromagnetic radiation. These interaction mechanisms include reflectance, emission, absorption, and scattering, and the proper blend of sensor wavelengths can reveal information about the types and condition of soil, vegetation, water, ice, and the atmosphere. One of the most popular regions of the electromagnetic spectrum employed by remote sensors are the visible bands that provide picture-like observations of the earth’s surface and the clouds that envelope the earth. However, many other portions of the electromagnetic spectrum are used to provide additional information about the earth and its resources. Infrared and microwave sensors provide infor- mation on atmospheric, surface, and even subsurface conditions. Microwave sensors have the added benefit of being able to observe the earth’s surface through clouds and the capability of operating day or night. 134 Introduction to Space Sciences and Spacecraft Applications

The combined data from passive and active sensors, spanning the elec- tromagnetic spectrum, allows users to study a host of important ecologi- cal, economic, and environmental subjects including crop coverage, health, and yield; land use and erosion; ocean and coastal conditions; and the effects of man on his environment in the form of pollution or deple- tion of the amount of ozone in the upper atmosphere.

Applications and Sponsorship

The genesis of remote sensing missions is usually associated with some applications goal that is sponsored by a civil governmental, military, or commercial organization. First efforts are usually exploratory research or proof-of-concept systems used to find the right sensor combinations and orbits for observation. If these research efforts prove successful, and a sustaining need can be established for the information that can be provid- ed by the space-based remote sensing system, then a decision may be made to continue with a series of operational satellites that provide a con- tinuum of information. For example, the aforementioned Tiros experi- ments led to the NOAA (National Oceanic and Atmospheric Administra- tion), GOES (Geostationary Operational Environmental Satellite), and military DMSP (Defense Meteorological Satellite Program) satellite series. This evolutionary process has led to a number of established oper- ational missions that continuously collect information about the earth, bal- anced by one-of-a-kind exploratory or scientific missions that constantly probe our earth, solar system, and universe in an effort to better explain the physical processes that govern our existence. Any attempt to rigidly classify missions erodes under the progress of technology, budgets, and sovereign or national interests. At present, most developed nations have the capability to create, operate, and sustain a wide-ranging space remote sensing program that can be segmented into intelligence, military, civil governmental, and commercial operational programs as well as a varied set of progressive scientific exploration pro- grams. With the exception of the protective and narrowly sovereign mili- tary and intelligence programs, many of the remaining mission categories are becoming shared efforts, with many nations combining their resources and technologies to produce cooperative benefits. The ever-expanding base of sensor technologies, missions, and international sponsors has pro- duced a robust interest in, and reliance on, space remote sensing-all stemming from the fact that the vantage of space allows man to look at Remote Sensing 135 himself and his world in a global sense for the first time in history. In order to continue, it is necessary to describe the physics and processes employed by space remote sensing systems to gain more detailed infor- mation of interest to users on earth.

Remote sensing is the collection of information about an object without being in physical contact with the object. Space-based remote sensors are also used for reference and control functions such as providing input for spacecraft attitude control, weapons guidance, spacecraft planetary land- ing, and spacecraft docking. In the context of this chapter, remote sensing will be restricted to data collection and to methods that employ electro- magnetic energy as the means of detecting and measuring target charac- teristics. This definition excludes electrical, magnetic, and gravity surveys that measure force fields rather than electromagnetic radiation. Remote sensing systems employ the electromagnetic spectrum in unique ways to send and/or collect natural or man-induced signals. The use of par- ticular wavelengths and signals for remote sensing can be based on: Natural radiation associated with the phenomena to be observed. Response of an object or material to an active signal created and trans- mitted to the target area by the remote sensor. The use of atmospheric or ionospheric interaction in the vertical path from the satellite to the earth, creating a class of sensors identified as sounders. The avoidance of atmospheric or ionospheric interference. Earth viewing systems concentrate on the microwave, infrared, and vis- ible regions of the spectrum from which information can be derived for a variety of specific imaging categories, including land mapping, specific target surveillance, and environmental monitoring. In addition to these more popular and commonly identifiable applications, space remote sen- sors yield important information on basic geophysical processes such as atmospheric temperature, water vapor, and liquid water; ocean surface wave height, length, and dynamics; surface wind speed and direction; sur- face temperature and ocean color; land area spectral evaluations; topo- graphic variations; and object or geological classifications. If the sensor relies primarily on an ability to discern the spatial charac- teristics of the objects in the field of interest, it is referred to as an hugex 136 Introduction to Space Sciences and Spacecraft Applications

Many sensors make primary use of the spectral characteristics of the response or the strength of the response to determine the context and con- tent of the measurement and are referred to as spectrometers or rudiome- ters. Whatever the primary method or measurement, all sensors have per- formance values described in terms of spectral (frequency-related), spatial (geometric), and radiometric (radiative-properties) dimensions. Moving or scanning sensors also provide a temporal (time-based) pattern of obser- vation, sometimes referred to as the revisit interval. Current research and development activities lead the way to eventual space-based evaluation of ocean and land subsurface physical and biological characteristics includ- ing soil moisture and bathymetric measurements. Finer spectral and spa- tial observations of land areas are on the horizon, and even surface pres- sure monitoring from space appears to be possible in the future.

Classification of Remote Sensors

Earlier in this chapter, we identified passive remote sensors that collect electromagnetic energy. The camera (without lighting aids such as a flash unit) is the classic example of a passive sensor: Incoming light impinges on sensitive film which retains the . Radiometers are devices that replace the film with detectors that are sensitive to the incoming radiation (which can be in other portions of the electromagnetic spectrum besides the visible frequencies) and convert the information to electronic signals that can be transmitted to and evaluated by users on the ground. Collec- tion devices include lenses and antennas that focus the energy on detec- tors such as solid state materials, charge coupled devices (CCDs), and other spectrally unique and energy-sensitive conversion elements. Sensors that create their own source of target or scene illumination are referred to as active sensors. These systems create and transmit energy that propagates from the satellite to the target, interacts with the target, and returns to the sensor. Active sensors include radars (altimeters, scat- terometers, real aperture radar (RAR) imagers, synthetic aperture radar (SAR) imagers) and lasers (laser altimeters, scatterometers, and imaging devices). These instruments have the benefit of being able to control the timelfrequency content, direction, and energy level of the illumination. Among other things, this characteristic permits active sensors to greatly improve on the spatial (geometric) performance of the system using meth- ods which will be described later. Remote Sensing 137

Geometry of Remote Sensing

The basic geometrical relationships between a satellite and the earth were described in Chapter 2. These are reproduced, along with some of the geometry specific to remote sensing, in Figure 6- 1. The right side of the figure displays the angles described in Chapter 2 associated with the angular field of view (Q) which represents the maxi- mum angular distance visible from the satellite when the remote sensor’s view is tangential to the earth’s surface. The maximum field of view (FOV) can be described as the total curved earth area visible from horizon to horizon by an observation in space, associated with the swath width distance described in Chapter 2. In most cases, a remote sensor concen- trates its instantaneous observation on a small portion of the total field of view using a lens, antenna, or some similar method of focusing incoming energy. This smaller area observed by the sensor at a particular time is known as the instantaneousfield ufview (IFOV) and will be described in more detail shortly. The geometry associated with remote sensing is shown on the left side of Figure 6- 1 ; these angles differ from those discussed in Chapter 2 in that they describe the relationship between the spacecraft and a point being viewed (the “target”). From the aspect of the satellite, the angle labeled 8, is called the bok (or nadir) angle and the complement of this angle is the depression angle (a). The distance from the sensor to the target R is known as the slant

Figure 6-1. Remote sensing geometry. Some particular geometrical terms are used to describe the relationship between a remote sensor and the earth or its target. 138 Introduction to Space Sciences and Spacecraft Applications range and is measured along the direction in which the sensor is looking (boresight). From the aspect of the point on the earth being viewed, the angle fiom the horizon to the satellite along the boresight is called the elevation (or grazing) angle (w),and the complement of this angle is called the incidence ang2e (ei).The angular distance from the nadir to the point being viewed, as measured from the center of the earth, is called the earth angle (y). Given the altitude of the satellite (h) and any one of the viewing angles, the slant range and all of the other angles can be calculated using:

R2 = (h + RJ2 + RZ - 2R, (h + k)COS($ (6- 1) and the law of sines: -=e=R R h+R, sin(y) si@,) sin(yl+ 90) Because the sensor may only be looking at a small portion of the available field of view at any particular time, a description of the area in view and a method of viewing the total area available must be developed.

Instantaneous Field of View. Just as a telescope of increasing power sees a smaller area on the face of the moon (for example), many remote sen- sors concentrate on a small portion of the FOV at any particular time. This smaller portion of the FOV visible by the sensor is known as the instanta- neous field of view (IFOV). The remote sensor’s ability to resolve spatial targets within the IFOV depends on the spatial acuity or dimensions of the detector’s picture element (pixel) or resolution cell which is a characteris- tic of the physical make-up of the sensing device itself. A passive remote sensor’s IFOV is dependent upon the wavelength scaled dimension of the collection or focusing element (antenna or lens) and the distance from the remote sensor to the target area being observed, as shown by (equation 6-3), sometimes referred to as the diffruction limit of spatial performance.

rFov=(;). (6 - 3)

where: h = primary wavelength of observation D = dimension of the aperture (lens, antenna, mirror, etc.) R = range from the sensor to the target area Remote Sensing 139

In equation 6-3, the term 3i/D is the sensor’s angular IFOV expressed in radians. This angular field of view, projected at range R, produces an IFOV covering a distance equal to the angular IFOV multiplied by R. A two- dimensional aperture produces a two-dimensional pattern, as exemplified in Figure 6-2 which also shows the pattern projected onto the ground. The antenna used in the figure is not symmetrical and, hence, does not produce a symmetrical pattern on the earth. Notice the inverse relationship between the IFOV and the antenna dimensions, as expected from equation 6-3. In the example shown, the antenna beam is pointed normal to the direc- tion of satellite motion. Translation of the sensor IFOV onto the ground plane is known as the ground instantaneous field of view (GIFOV). Description of the GIFOV in the direction that the beam is pointing (referred to as the range or cruss-truck direction) involves simple trigonometry as shown in equation 6-4. hR GIFOVRANGE= (6 - 4) D, shy, In the direction that the satellite platform is translating, referred to as the azimuth or alongtruck direction, the GIFOV is simply equal to the IFOV, as shown in equation 6-5.

IMUM DIRECTION

PUSHBROOM

RANGE DIRECTION Figure 6-2. Instantaneous FOV and scanning. The field of view of the sensor can be superimposed upon the ground and scanned in several ways. 140 Introduction to Space Sciences and Spacecraft Applications

Sensor Scanning. In most cases, a remote sensor would ideally cover as much of the observable area (field of view) as possible. To do this with an IFOV smaller than the FOV, the sensor must shift its concentration in some way, as a function of time, over the FOV. One important point to keep in mind is that the satellite (in all cases except for a geostationary satellite platform) is itself in motion over the planetary surface. There are many methods that can be used to change the direction of the sensor’s IFOV, including mechanically moving a telescope, mirror, or anten- na; electronically shifting an antenna beam or optical path; and/or arranging a line or array of sensors with respect to the satellite’s orbital motion. The scan can also follow a variety of patterns, some of which are illustrated in Figure 6-2. Many sensors scan from side to side across (normal to) the satel- lite nadir track. This type of pattern is known as a cross-truck scan. Since the grazing angle and range are constantly changing as the scan gets farther from nadir, the GIFOV also constantly changes in the range direction. Some sensors, such as microwave radiometers, hold a constant nadir (look) angle and then scan the sensor around the satellite nadir in a circu- lar pattern known as a conical scan. Or a sensor can be held stationary and simply use the satellite motion to move the field of view across the sur- face in a technique called a pushbroom scan. Several modern optical sen- sors employ thousands of cross-track detectors in the sensor focal plane, and use the satellite’s orbital translation to create such a scan. These and other variations on these themes are used to permit the selective IFOV of the sensor (representing the limit of a sensor’s spatial acuity) to collect information over a broader range of the satellite field of view. If a contiguous scan coverage is required, the scan rate (time for the sensor to complete one full sweep of the desired areas) is set so as to allow each scanned area to overlap slightly with previously scanned areas so that no spots are missed. Many sensors do not require a contiguous scan and instead may use a sampling scan to achieve their measurement goals.

Example Problem:

A remote sensing satellite is in a circular orbit at an altitude of 1,OOO km. The sensor is a radiometer with its beam aligned normal to the Remote Sensing 141

satellite direction of flight and held at a fixed look angle of 30 degrees. This orientation is identical to the example shown in the previous fig- ure. The aperture dimensions are 5 m (Dazimufh)x 1 m (Dcmss.uack)and the radiometer center frequency is 2 GHz. If the radiometer is held at a fixed look angle of 30", determine the slant range to the ground, IFOV, and GIFOV for the along-track and cross-track dimensions.

Solution:

Using equation 6-2 (knowing the look angle, h and RJ we can solve for all the other unknowns:

(w + 90) = 35.3" or 144.7" (must be > 90')

3i = 5-34"

R = 1186.8 km

From the center frequency of observation:

IFOV,,s.,k = 178.1 km

GIFOV,,,,~,,,k = 218.2 km

Observation of Electromagnetic Radiations

The electromagnetic spectrum was described in Chapter 4 along with the relative frequencies and energies radiated by the sun and the earth due to their respective surface temperatures. In the case of passive sensors, these natural radiations, and their interaction with the earth and atmosphere, are what is sensed. Active sensors operate in exactly the same manner except that they provide a known spectral illumination of the target area and then sense the reaction of the target and atmosphere to the transmitted frequencies.

Target Spectral Interaction Mechanisms. A number of interactions are possible when electromagnetic energy encounters a target or scene. The 142 Introduction to Space Sciences and Spacecraft Applications interactions can take place in a plane or two-dimensional surface (su~uce phenomena) or in three dimensions, resulting in interactions called vol- ume phenomena. These surface and volume interactions can produce a number of changes in the incident electromagnetic radiation, such as changes in magnitude, direction, wavelength, polarization, and/or phase. A remote sensor measures and records these changes, and the resulting data are interpreted to identify the characteristics of the target or scene that produced the changes in the electromagnetic radiation. During the interaction between electromagnetic radiation and matter, both mass and energy are conserved according to basic physical princi- ples. The following complex interactions may occur:

Radiation may be transmitted, that is, passed through. The velocity of the electromagnetic wave may change as it is transmitted from a vac- uum into other mediums. This velocity change is characterized by the medium’s index ofrefraction (n) which relates the velocity of electro- magnetic propagation in the medium (c,) to that in a vacuum (c) as:

n = CJC, (6-6) Radiation may be absorbed and give up its energy, largely due to heat- ing the substance or medium. Radiation may be emitted by elements in a scene as a function of ele- ment structure and temperature. As we saw in Chapter 4, all matter at temperatures above absolute zero emits energy. Radiation may be scattered. The scattering process may be preferential, tending to propagate in some specific directions, or it could scatter in all directions and be lost ultimately to absorption or further scattering out- side of the sensor FOV. If the illumination source is the sensor itself, the energy returned to the sensor by this process is called backscatter Radiation may be reflected, unchanged, from the surface of a substance with the angle of reflection equal and opposite to the angle of incidence.

These interactions with any particular target or form of matter are selec- tive with regard to the wavelength of electromagnetic radiation and are spe- cific for that form of matter, depending primarily upon its optical orientation with the sensor boresight, its wavelength scale-scattering surface or volume properties, and its atomic and molecular structure. These interactions between matter and energy provide the spectral and radiometric basis for remote sensing. Remote Sensing 143

Radiative Transfer Equation. The amount of energy that a remote sensor receives from a target area is very much frequency-dependent and may come from a variety of sources as a result of the interactions described above. Figure 6-3 demonstrates the types of energies that a remote sensor may be looking to receive and reveals some of the interactions that energy may have during its transmission through the atmosphere. In the figure, IT represents the total energy received by the satellite sen- sor which has been designed to observe phenomena within a specific fre- quency range. In the situation depicted, the illumination of the scene is provided by the energy of the sun which has the known spectral charac- teristics described in Chapter 4. In other cases, the received energy may instead be produced by the natural radiations of the earth and atmosphere, or may be supplied by the transmitter of an active remote sensor. Each of the radiations which make up the total received energy is also a function of frequency and, for some observed frequencies, some of the components shown may not contribute to the received energy at all. As shown in the figure, the sun's radiation, denoted by F(h),illuminates the scene from an angle 8, called the zenith angle. One source of energy received by the remote sensor may come from some of this radiation that is scattered off small particles within the atmosphere. This aerosol scut- tered energy (IA)is uniformly radiated from the particles in all directions. Another type of scattering of energy radiates in an anisotropic (not uni- form) manner from the molecular constituents of the atmosphere. This

-"a- \ 'R" '0, \ yI\

Figure 6-3. Remotely sensed energies. The energies received by a sensor in space may come from a number of sources and/or interactions. 144 Introduction to Space Sciences and Spacecraft Applications

RayZeigh scattering (pictured as a peanut-shaped pattern in the figure) is very direction-dependent and the scattering angles (denoted by a+ and a-) are a function of both the sun’s zenith angle and the satellite look angle. The total amount of Rayleigh scattered energy received by the remote sensor may come from different sources: that scattered from direct illumination by the sun (Ih+), the reflection of this scattered energy off the surface of the earth (I%-), and that portion of the sun’s energy that first reflects off the surface of the earth and then gets Rayleigh scattered (1%). Finally, the satellite sensor may receive energy directly from the earth’s surface. The energy level of the radiation leaving the surface (whether the earth’s natural temperature-related radiations or reflection of solar or man-induced signals) is denoted by I,,. Transmissivity. Since the energy received by a remote sensor must prop- agate through the atmosphere, this energy may be changed during transit due to the different scattering modes and the absorption and reradiation of energy by the components of the atmosphere, all of which are highly fre- quency-dependent. The effect of the atmosphere on electromagnetic prop- agation is modeled by a transmissivity factor denoted by ~(1).Figure 6-4 shows the transmissivity of radiation through the atmosphere as a function of percent transmission versus wavelength. The figure also indicates the atmospheric constituents responsible for the poor transmittance of some frequencies through the absorption mech- anism described earlier in Chapter 4 using a similar curve showing the spectral energy propagation through the atmosphere. Figure 6-4 only

100 Uw Q != 60 2 k

6B 20 e

12 3 4 5 6 7 8 9 10 11 12 1314

WAVELENGTH (pm) Figure 6-4. Atmospheric transmissivity. Some atmospheric constituents react with certain wavelengths, absorbing some or all of the associated energy. Remote Sensing 145 shows a portion of the infrared spectral band and not the entire electro- magnetic spectrum. Frequency ranges with poor transmissivity are known as absorption bands, while regions where radiations pass through rela- tively unaffected are called windows. As an example, we can see well dur- ing the day because the atmosphere is practically transparent to the visi- ble frequency range; but a couple of hours at the beach doesn’t burn us too badly because most of the ultraviolet radiations of the sun are absorbed high in the atmosphere. A particular remote sensor’s purpose may be to observe and measure some phenomenon at the earth’s surface, in which case it must negotiate frequencies with significant atmospheric absorption and use the window frequencies. Many sensors, sometimes called sounders, wish to measure atmospheric or ionospheric conditions and will intentionally select frequencies of observation that interact with and mea- sure high-altitude phenomena. The component of the total energy received by the remote sensor from just the su$ace is given by:

where

which shows the dependency of the received energy on the spectral output of the sun F,) and the effects of the transmissivity z(k) on the energy prop- agation. Similar equations exist for the other components of the total received energy which are also dependent on these factors. The total ener- gy received by the remote sensor is the sum of these radiations which, when expressed in equation form, constitute the radiative transfer equation:

where the total Rayleigh scattered input is given by IqtOt,,and I* is the received aerosol scattered component. The reduction of this raw received information is quite complicated and requires some prior knowledge of the behavior of the atmosphere. Over many years, algorithms have been developed which are used to reduce the unrefined sensor source or raw data and identify the physical processes which contribute to the total energy received by the remote sensor. Algo- 146 Introduction to Space Sciences and Spacecraft Applications rithm development is based on a theoretical understanding of the physical processes involved, bolstered, and refined by empirical experiments using aircraft and satellite remote sensor observations. From this, the basic geo- physical properties discussed earlier, such as temperatures, atmospheric constituents, and land resources, can be distilled.

MEASUREMENT DIMENSIONS OF REMOTE SENSORS

All remote sensors observe in the context of specific spectral regions, and within each spectral band, a sensor further defines an observation in terms of space, time, and signal strength. The dimensional extent of a sen- sor’s ability to observe has been defined as the swath width or field of view (FOV). Within the FOV, a sensor’s spatial acuity is defined within the IFOV elements (pixels). Each IFOV element is further defined in terms of the strength of the signal identified with a specific observation or scene location. Considering all of the IFOV elements that combine to con- stitute a scene, the variation in signal strength produces contrast. Sensors depend on a satellite’s orbit and the sensor’s scan pattern to determine the precise observation time and sequence for a particular earth location.

Spatial Performance

Passive Sensor Spatial Resolution. The spatial performance of passive sen- sors is defined by the wavelength scale dimension of the focusing aperture (in radians) and the range between the observing sensor and the observation area. The satellite orbit and sensor scan pattern can alter the exact dimensions of the earth-projected sensor IFOV, but the beam projection is fundamental- ly diffraction limited. Active sensors can go a step further in gained improved resolution in both the range and along-track dimensions.

Active Sensor Spatial Resolution. Active sensors can improve on the resolution in the range or beam direction by controlling the length of the pulse of energy transmitted by the sensor. This is known as pulse ranging and is shown in Figure 6-5. The active sensor sends out a pulse of energy with a pulse length of z seconds. This pulse propagates in the direction of the beam (boresight) and illuminates the IFOV area, after which some of the energy reflects back to the receiver. In this case, the range resolution (R,) depends on the pulse length and the signal propagation velocity (c,) which is approxi- Remote Sensing 147

DEPRESSION ANGLE

TARGETS A&8 TARGETS C&D

Figure 6-5. Active sensor range resolution. Active sensors can improve on resolution in the range direction by controlling the pulse length of the transmitted energy. mately the speed of light for most active sensor frequencies and space-to- earth viewing scenarios. The relationship is given in equation 6-8.

This sensor resolution is achieved in the slant range plane and must be translated into the ground plane as a function of the local target incident angle (ei)as shown in equation 6-9.

Radars and lasers improve their spatial resolution in the range or beam direction by using this technique. Active radar sensors can either stop at this point or continue this spatial resolution improvement process by devising a method of improving the along-track or azimuth spatial performance of the sensor. The two basic types of active imaging radar systems are real-aper- ture radar (RAR)and synthetic-aperture radar (SAR), which differ pri- marily in the method of achieving a desired resolution in the azimuth direc- tion. The RAR obtains its along-track resolution through the wavelength scale physical dimensions of its antenna in the along-track direction. As a result, the RAR uses an antenna of the maximum practical length to produce a narrow, diffraction-limited, angular beam width in the azimuth direction. 140 Introduction to Space Sciences and Spacecraft Applications

Synthetic-aperture radar systems employ a smaller antenna that transmits a relatively broad azimuth or along-track beam as shown in Figure 6-6. A SAR “synthetically” obtains its along-track resolution by moving the real antenna beam in relation to the target. In the most basic sense, there must be a translation of either the target through the real beam, the real beam through the target, or a combination of both processes. As shown in the figure, as‘the radar beam translates the target, the radar sends a string of carefully prescribed pulses illuminating and receiving backscatter responses from the target area. The SAR preserves and saves the ampli- tude and phase histories of each of these responses. A special image sig- nal processing system performs a target specific weighting, shifting, and summing process focusing on one resolution element at a time and, as a result, achieves an along-track resolution that is independent of range and wavelength, becoming equal to one-half of the actual antenna length in the along-track direction, as shown in the figure.

--au -QAT 2au 2

Figure 6-6. Aperture synthesis. A synthetic aperture radar (SAR) uses signal processing over many radar returns to obtain resolutions comparable with normal radars of much larger apertures. Remote Sensing 149

Spectral (Radiometric) Performance

The spectral performance of a sensor depends on the frequency (or wave- length) of the radiation which the sensor is designed to detect. The wave- length(s) used to make a measurement is selected based on the radiative (radiometric) properties of the target and/or on other considerations, such as interaction with or avoidance of the atmosphere or ionosphere. For instance, in clouds, water occurs as aerosol-sized particles of liquid rather than vapor. Clouds absorb and scatter electromagnetic radiation at wave- lengths less than about 0.3 cm. Sensors are designed to take advantage of this effect to measure atmospheric characteristics. Only radiation of microwave and longer wavelengths is capable of penetrating clouds without being significantly scattered, reflected, or absorbed. These (radar) fkequen- cies are commonly used when information below cloud layers is desired. The spectral performance of a system is a measure of how well the sys- tem uses these frequencies (or combinations of frequency ranges) to extract information on a desired target. Much of the success of a remote sensor depends on knowledge of the radiometric properties of the target. These targethcene characteristics determine the relative strength or inten- sity of observed signals based on their reflectance, absorption, or other properties for the particular frequencies involved.

Temporal Performance

The temporal performance of a remote sensor is a function of the repeated observations of a phenomenon based on the expected rate of changes in that phenomenon. Global winds change on an hourly or daily basis while changes in the earth’s geodetic shape or magnetic field may take years or millennia to occur. The observation rate of a specific phe- nomenon or area is a function of the orbit in which the satellite is placed and the sensor FOV produced by the scan pattern. An orbit and scan pat- tern can be preferentially designed to create a revisit sequence and local time observation tailored to a particular application.

Early remote sensing satellites were technology and lift capacity con- strained, resulting in relatively simple sensors designed to observe singular regions of the electromagnetic spectrum. Modem satellites tend to maximize 150 Introduction to Space Sciences and Spacecraft Applications in-orbit returns by investigating several electromagnetic regions simultane- ously. Figure 6-7 shows the frequency bands used by modem remote sens- ing systems, with the regions of interest aligned with particular frequency ranges. Note that radiation in the short wavelength portion of the lR band and the long wavelength portion of the UV band is detectable by photographic film. Therefore, these wavelengths are included with the visible wavelengths and are designated as the photographic remote sensing band. In conclusion, it should be noted that mission application objectives and remote sensors drive the spacecraft and orbit design. The sensor systems are designed first to give the desired performances as described above, and the other spacecraft systems (such as attitude control, power, and commu- nications, which will be discussed in a separate chapter) are created to sup port these sensors in the collection and transmission of the observed data to the ground for synthesis, geophysical analysis, and dissemination.

REMOTE SENSING BANDS

PHOTO THERMAL GRAPHIC IR RADAR -I -

I SPECTRAL GAMMA BANDS X-RAY [ 'EF MICROWAVE RADIO E

I I I II II Ill .003 .03 .3 3 30 .3 3 30 300 .3 3 30 3 30 300 om nm nm nm nm pm pm pm pm an cm an m m m

WAVELENGTH Figure 6-7. Remote sensing bands. Modern remote sensing satellites may contain multiple sensors designed to investigate several areas within remote sensing bands simultaneously.

REFERENCES/ADDITIONAL READING

Cantafio, L., Space-Based Radar Handbook. Boston: Artech House, 1989. Sabins, E, Remote Sensing Principles and Intelpretation. 2nd ed. New York: W.H. Freeman and Company, 1985. Schnapf, A., Monitoring Earth's Ocean, Land, and Atmosphere from Space- Sensors, Systems and Applications. Washington, D.C. : Ameri- can Institute of Aeronautics and Astronautics, 1985. Remote Sensing 151

Manual of Remote Sensing, Volumes Z and ZZ, 2nd ed. Washington, D.C.: American Society of Photogrammetry, 1985. Beal, R., Spaceborne Synthetic Aperture Radar for Oceanography. Balti- more: The Johns Hopkins University Press, 1981.

1. A sensor is to be placed into geostationary orbit to monitor its field of view constantly for infrared signals, indicating possible ICBM launch- es. The sensor is to consist of a 10 x 10 pixel square array of photode- tectors looking at the frequencies associated with temperatures between 1,500 and 3,000" K (1.29 pm median wavelength). The electronic out- put of the photodetectors is a sinusoid with a changing amplitude that represents the detected temperature. This signal is to be digitized using 16 bits per word with a sampling frequency of 1 KHz. Draw a sketch and determine the geometry of the situation, including look angle (en),earth angle (y), and elevation angle (v),for the maximum slant range (R-). (Hint: v = 0 at Rmm.)Determine the size of each pho- todetector required to give a ground resolution (the JFOV for one detec- tor) of 10 km when looking straight down. What is the difference between this resolution and that when the sensor is looking out at Gax?Compute the system IFOV and GIFOV at nadir and at R-. Also, compute the tem- perature sensitivity and transmission rate of the system. Finally, discuss how you would attempt to scan the entire visible earth's surface for com- plete coverage. Based on your scan method and the given sampling rate (1 sample = 1 area covered), estimate how long it would take your system to scan the entire visible surface once (assuming no overlap between areas scanned each sample). 2. Determine the size aperture a remote sensor would require if it were looking for possible ICBM temperatures between 1,500 and 3,000 OK (1.29 pm median wavelength) and wished to see the entire earth's sur- face possible at all times from its geostationary position while always looking straight down at nadir. Use Figure 6-1 to work out the geome- try of the problem. Hint: To do this you need to find the distance per- pendicular from the nadir line (between the satellite and the center of the earth) and the point on the horizon of the satellite's field of view. 152 Introduction to Space Sciences and Spacecraft Applications

3. If the sensor in Exercise 2 is a 10 x 10 pixel array of CCDs generating byte-sized words each sample, and the transmission rate is limited to 200 Mbps, determine the sensitivity of the system and the maximum sampling frequency for the system. 4. A photographic reconnaissance satellite operates at an altitude of 150 km, always looking straight down. It has a lens with a diameter of 0.5 meters which projects the received images onto a 64 x 64 pixel array of CCDs with a receiving capability from 5 x 1014 Hz to 6 x 1014 Hz (0.69 to 0.59 in wavelength where lfi = 1 micron = meter). Pulse code modulation (PCM) is used to digitize the data, but the system is limit- ed to 16 bits of information for each CCD each sample. Fifteen thou- sand pictures are taken each second (for overlap). Determine the fol- lowing: a. The field of view of the image. Use the median wavelength of 0.555 for the calculation. b. The number of quantization levels available. c. The frequency sensitivity of the system. d. Total number of bits produced each sample. e. Transmission rate in Mbps. f. (EXTRA CREDIT) If the CCD array is 1 square meter, what is the image resolution of the system? (OR) Determine if the picture-taking frequency is sufficient for image overlap. 5. A remote sensor has a 32 x 32 pixel array of infrared-sensitive charged- couple devices (CCDs). The range of sensitivity for the CCDs is centered between 250 OK and 500 OK. Pulse code modulation (PCM) is used to digitize the data, and the sensitivity range has been divided into 16 quan- tization levels. Sampling fiequency is 10 Khz. a. What is the temperature sensitivity (represented by each Q-level)? b. How many bits of information does a CCD produce each sample? c. What is the total number of bits produced each sample by the array? d. What is the transmission rate of the system in bits per second (bps)? 6. A passive remote sensor has a 1 meter diameter circular lens and oper- ates at a frequency of 3 x lOI4 Hz. Calculate the IFOV at a range of 1,OOO km. What would be the ground IFOV if the grazing angle was 60°? Remote Sensing 153

7. A remote sensing satellite is operating at an altitude of 800 km. The maximum nadir look angle of the scanning sensor is 30" side to side rel- ative to the satellite's ground track. Calculate (a) the grazing angle and (b) the slant range corresponding to the maximum look angle, and (c) the total horizon-to-horizon FOV distance in km. 8. An active radar sensor sends a pulse of 3 x sec (3 nanosec) which strikes the earth at a 30" incidence angle. Calculate the ground resolu- tion in meters of the radar. CHAPTER 7 Satellite Navigation

The concept of using satellites for navigation dates back to Sputnik 1 in 1958 when engineers monitoring this first orbiting body observed that the substantial Doppler frequency shift of the received telemetry signal could be used to determine accurately the satellite’s position and orbit. It was quickly shown that the inverse situation was also possible; that if the satel- lite position (orbit) were known accurately, an observer’s location on the earth could be determined. After several years of work at the Applied Physics Laboratory (APL) of Johns Hopkins University, a prototype satel- lite was developed and launched, and satellite navigation was born. The first operational systems were part of the Navy Navigation Satel- lite System known as “Transit.” The system uses the Doppler positioning method described further in this chapter and requires information gathered from only a single satellite, over the time of passage overhead, to produce a fix. The newest generation of navigation satellites are part of the Glob- al Positioning Satellite (GPS) system. This system uses a pulse ranging method of positioning and requires four satellites to be in view of the receiver simultaneously for a complete solution. Though the requirements are more stringent, GPS provides real-time, three-dimensional position, track, and speed with accuracies many times better than the earlier Tran- sit system. POSITION DETERMINATION USING DOPPLER TECHNIQUES

The original method of determining position from satellite signals is through observation of the Doppler shift in frequency of a known signal as the satellite passes over the observation area. The shift in frequency is a unique function of the motion of the satellite relative to the observer’s position. To compute an observer’s location accurately, the position and orbit of the satellite and the frequency of the transmitted signal must be known. Also, timing between the transmitting satellite and the observer’s

154 Satellite Navigation 155 receiver is important. Most of these requirements are provided by a mes- sage contained in a separate signal transmitted by the satellite (the navi- gation message),which reports the satellite ephemeris (position and orbit information) and a timing signal. The frequency of transmission is known by all users and must be provided by a very stable source aboard the satel- lite. The following sections show how this information is combined to allow an observer to compute a position on the earth.

Doppler Ranging

The frequency received from the satellite (fR)consists of the transmit- ted frequency (fT) plus a Doppler frequency (+,) due to the relative motion between the satellite and receiver. The receiver also has a stable oscillator which produces a reference signal at approximately the same frequency as transmitted by the satellite (fG).Subtracting this frequency from the received frequency, the receiver determines the shift in frequen- cy in terms of a Doppler count. The Doppler count represents a counting of the number of frequency cycles occurring between timing signals trans- mitted by the satellite. In equation form:

where tl + Rl/c represents the time the signal was received after it was trans- mitted from the satellite at time tl and traveled to the receiver over a slant range of R1at the speed of light c. Expanding equation 7-1 into two parts:

t2 +R~lc t2 +R2 IC NI=j f,dt-! tl+Rl/c fRdt (7 - 2) ti +Rl/c

The first part is simply the integral of a constant since the receiver pro- duces a stable frequency (fc),but the second part contains the changing received frequency fW However, the second integral also represents the number of cycles received between the two timing signals (sent by the satellite) which must be equal to the number of cycles sent by the satel- lite between those same two times. This allows us to rewrite the second part in terms of simply the transmitted frequency and the transmitter tim- ing signals: 156 Introduction to Space Sciences and Spacecraft Applications

t2 +R2/c N, = fG dt - jt2fT dt (7 - 3) ti +Rl/c 11

Because fT is also considered constant and known, both the above inte- grals can be performed resulting in:

N, =f, [(t2 +%)-(tl +%)]-fT (t2 -tl) (7-4)

By combining terms, equation 7-4 can be rewritten as:

The first term of equation 7-5 is just the difference between the receiv- er’s oscillator frequency and the satellite’s transmitted frequency, which is a constant term and can simply be ignored. The second term is a measure of the change in range between the satellite and the receiver during the time between two timing marks. This information is used to determine the receiver’s position. Notice that this range difference is given in terms of wavelengths of the receiver’s generated frequency (fc).

Computation of Position

Computation of position usually requires a small digital computer. The computer determines the satellite position at each timing signal from the ephemeris information transmitted by the satellite. The computer also uses an approximate position of the receiver, either inputted by the user or estimated from the last fix. The expected change in slant range between the satellite and the receiver (based on the estimated position) is comput- ed and compared to that actually measured. The estimated position is then shifted in such a way that the differences are decreased. Through several iterations of this procedure, the estimated position is refined until the dif- ferences are reduced to an acceptable value. As mentioned earlier, the Transit system requires only a single satellite to be in view to determine position. The iterative process is repeated for as long as the satellite is in view resulting in an increasing positioning accu- racy with time, as illustrated in Figure 7-1. If the receiver is in motion, the computer must know the motion accurately to compute the Doppler-based Satellite Navigation 157

Figure 7-1. Doppler positioning. Successive position iterations may be achieved from a single pass of a single Transit satellite. slant range change precisely and to know the change in receiver position over the observation time. Multiple fixes can be combined to provide posi- tion, track, and speed for subsequent iterations resulting in increased accu- racy. If a single satellite is used, a dead reckoning (DR) plot is used between times the satellite is below the horizon to give an estimated posi- tion, track, and speed for the next satellite encounter.

Errors and Accuracies

The theoretical accuracy of the Doppler ranging method is associated with the wavelength of the frequency used for the navigation signal. For instance, a 400 MHz signal (currently used by Transit) corresponds to a 0.75 meter wavelength and a similar theoretical accuracy (for a single fix). Unfortunately, many sources of errors exist which affect the accura- cy of satellite navigation systems. Refraction Errors. There are two sources of refraction errors as the satel- lite signal propagates between the satellite and the receiver. The first is introduced by the ionosphere which changes the path of propagation of the signal as described in Chapter 5. This change introduces a shift in the frequency of the signal which would affect the determination of the slant range change between the satellite and receiver. In order to compensate for this shift, navigation satellites transmit over two different frequencies and compare the Doppler measurements made at each to reduce the error introduced by the ionosphere. 158 Introduction to Space Sciences and Spacecraft Applications

The second source of refraction error is the troposphere which affects the speed of propagation of the signal as it travels through the atmosphere. This error changes with atmospheric changes, such as temperature and humidity, and with elevation angle between the receiver and the satellite (signals travel through more of the atmosphere, and are affected more, with lower elevation angles). The troposphere affects all frequencies sim- ilarly and is not as easily subtracted as the ionospheric effects. Receiver station compensation for tropospheric errors, if any, is usually done using simple atmospheric models.

Position Errors. The accuracy of the fix obtained from Doppler posi- tioning, with respect to ground latitude and longitude, is a function of the receiver’s position with respect to the satellite orbital plane and knowl- edge of the receiver’s altitude above the reference “spheroid” (the surface defined as if the earth were actually a perfect sphere) on which satellite position is based. (Note: The “geoid” is another common reference sur- face and represents the surface defined by mean sea level worldwide. Both the spheroid and the geoid are different from the actual topography of the earth. Knowing one’s altitude, in this sense, is not as simple as it may seem.) Figure 7-2 shows how the same range rate information would report a different position if the altitude were not accurately known. Notice that, for the same error in altitude, the error in computed position is greater if the receiver is closer to the satellite orbital plane. Additionally, as was mentioned earlier, receiver position, track, and speed are required for accurate positioning. Estimated position is not as important as track and speed, as this error will be eliminated through the iteration procedure of the computer’s correlation with slant range rates. However, track and speed are very important as they affect the computa- tion of the Doppler shift between satellite and receiver.

SURFACE I Figure 7-2. Altitude error. Computed position may have an error associated with uncertainty of “altitude” with respect to the reference plane. Satellite Navigation 159

Satellite Errors. Errors attributable to the satellite include rounding in the ephemeris data transmitted by the satellite to the receiver, and inaccura- cies of this reported information due to orbital parameters such as atmos- pheric or solar pressure drag and inaccuracies of prediction of the satellite orbit over time due to insufficient knowledge or modeling of the geopo- tential (gravitational) model of the earth. Satellite ephemeris is computed frequently by ground stations and the navigation message updated often to reduce these errors. Additionally, satellite and receiver clock and oscil- lator instabilities may introduce another, not insignificant error.

PULSE RANGING AND PHASE DIFFERENCE POSITIONING

Experience with the first Doppler-based satellite navigation systems pro- duced newer ideas for conducting positioning from orbiting satellites, which offered increased accuracies and eliminated some of the sources of errors affecting the earlier method. Doppler methods compare estimated range rates with those derived from the Doppler shift to determine a position correction. A simpler concept involves deriving one's actual range directly from the satellite transmitted signal. The idea is depicted in Figure 7-3. If a signal (a pulse, for instance, as shown in the figure) were sent out by a satellite at a known time, the exact range from the satellite to the receiver could be found from:

R=c,At (7-6) where c, is the speed of propagation of the signal and At is the time between transmission (kmit)and reception (t,,") of the signal. The range represents a sphere, somewhere on which the observer would lie. An exact position, in three dimensions, could be obtained by the intersection of

Figure 7-3.Pulse ranging. Distance (slant range) between a transmitter and receiver can be determined by knowing the time it takes a signal to travel. 160 Introduction to Space Sciences and Spacecraft Applications three of these range spheres (once again, knowing precisely the position of the transmitter[s]). The concept is quite simple, but actual implementation of the idea poses some real problems. For example, the signal sent by the satellite, from which the observer will compute range, must allow the receiver to determine exactly when a particular event was transmitted. Ranging Signal

Specifying the particular “event” to be used to compute range lends itself very well to digital techniques. To determine exactly which “event” has been received (assuming the satellite transmits continuously), a digi- tal code with a unique pattern could be used. The receiver picks up the sig- nal and compares it to the same code stored in memory to find where along the pattern the received information is located. A convenient method for impressing the digital code onto a suitable carrier wave is through biphase modulation (review Chapter 5 if neces- sary). The “event” is a change (or no change) of the phase of the carrier wave by 180” each time the mate of the digital code changes. An example of a biphase modulated signal is shown in Figure 7-4. The transmission rate of the digital sequence determines two important characteristics of the system. A small time between phase changes (Q allows the receiver to more accurately determine range from the transmit- ter, but a larger pulse period ensures better detection of the pulse and error- free use of the information. Tradeoffs between these objectives results in the modulation frequency used (which affects the receiver minimum signal-to- noise ratio) and also determines the theoretical accuracy of the system.

0 1 0 0 1 1 0 1 ’.. DIGITAL STRING

”BED

BI - PHASE MODULATED SIGNAL

Figure 7-4. Biphase modulated signal. An “unrepeating” code is modulated onto a carrier frequency to determine the timing “mark” for pulse ranging. Satellite Navigation 161

Timing Signal

Another difficulty posed by the ranging concept is determining exactly when the particular “event” was sent. Solving this problem involves two steps: correlation and synchronization.

Correlation. Figure 7-5 displays the three signals involved in pulse rang- ing. The first signal is the biphase modulated signal transmitted by the satellite Below it is the identical signal stored in memory of the receiver (SI($. Note that this stored signal may not be exactly synchro- nized in time with the transmitted signal. The third signal shown (S2(t))is the received signal (ignoring such distractions as noise). Both the replica signal and the received signal are digitally sampled and then the signals are compared using a digital cross correlation function of the form of equation 7-7:

t (Note: Summation in discrete-time [digital] mathematics is the same as inte- gration in conventional math.) Summation over many values oft results in a particular time (to)for which the summation is maximum. This represents the offset time, or time between the occurrence of the same particular event (bitlphase change) in the received signal and the stored replica code.

Synchronization. The offset time found above does not yet represent the true At between transmission and reception of the signal, as the satellite

Figure 7-5. Pulse ranging signals. The relationship between these signals is used to determine the times required for an accurate ranging. 162 Introduction to Space Sciences and Spacecraft Applications and receiver clocks may not be in perfect synchronization. This clock off- set time is shown in Figure 7-5 sketch as the tc time difference between the satellite signal and the receiver’s replica code. If it is assumed that the satellite is the absolute time reference, then this clock offset time would produce an error in range if not allowed for. The method for determining the value of the clock offset time is described as part of the ranging process described next.

Range Determination

The range equation (eq. 7-6) would now look like:

R = C, (tof t,) (7-8) In terms of an X, Y, Z coordinate system, this equation would be of the form:

R1 = I(X,(t,)- XI2 + (YI(t1) - YI2 + (Zl(t1)- z)211/2f cdtc) where the X1(,,,, Y1(,,),Zl(tl) terms represent the position of the transmit- ter (satellite) at the time of transmission (tl). X, Y, and Z represent the position of the receiver computed from the range information derived from the correlation offset time c, to, and c,,, tc represents the range error due to the satellitekeceiver clock offset time. Assuming that the position of the satellite is known precisely at all times (via a navigation message as described earlier), this equation has four unknowns involving the X, Y, Z position coordinates of the receiver and the receiver clock offset time tc. To solve for these unknowns, at least four range signals must be available at the same time to solve four simi- lar equations simultaneously:

R1 = [(X,,,,) - XI2 + (y~(q)- Y12 + (Zl(tl) - Z)211” * Cm(tcI

R2 = [(X2(t2) - XI2 + (Y2(t2, - Y12 + (G(t.7) - Z)21’” cm(tc>

R3 = [(X3(t3) - XI2 + (Y3(t3) - V2+ (Z3(t3) - Z)21’n * cm(tc)

R4 = [(X4(,, - X)’ + (Y4(t4) - Y)’ + (Z4(t4, - Z)21’’2 * Cm(Q

Note that the same receiver clock offset time tc appears for each of the received signals. This assumes that each of the received signals comes from satellite transmitters synchronized exactly with absolute time. If this were Satellite Navigation 163 the case, then even if the above ranges were computed from four different satellites, when the information was received all the times would be the same, i.e., tl = t2 = t3 = t4. In practice, another time correction (included in the navigation message) is incorporated to synchronize the transmitters on board the different satellites themselves to an absolute reference time.

Errors and Accuracies

The refraction-type errors described earlier affect all types of transmis- sions through the atmosphere and decrease the accuracy of the ranging process as well. As with Transit, GPS satellites transmit on two different frequencies to compensate for ionospheric effects on the signals. These much higher frequencies (1.58 and 1.23 GHz) are also less affected by the troposphere and additionally allow a much higher bit rate resulting in greater accuracies. No estimated position or track and speed information is required because solution of the four range equations simultaneously gives a very accurate three-dimensional fix. Satellite errors can be minimized with more accurate and stable oscillators, better geopotential models and more stable orbits, and updating of the navigation message more frequently. Table 7-1 compares the accuracy of some of the more common naviga- tion systems. TACAN (Tactical Air Navigation), Omega, and Loran are examples of existing ground-based radio navigation aids.

Table 7-1 Accuracy of Some Common Navigational Systems

~~ ~~ ~ Position Velocity Accuracy Accuracy System (meters) (m/sec) Comments

~~ TACAN 400 none Line of sight air navigation. Omega 2,200 none Worldwide radio navigation. Loran-C 180 none Localized area (U.S. shore) radio navigation. Transit 200 none Worldwide, but up to 100 min. between satellite passes. GPS 1.5" 0.1* Global, 24-hr, all-weather availability.

*Accuracies only available to authorized (military) users. Public accuracies are appmximarely a magnitude less. 164 Zntmduction to Space Sciences and Spacecraft Applications

REFERENCEWADDITIONAL READING

Logsdon, T., The Navstar Global Positioning System. New York: Van Nostrand Reinhold, 1992. Dutton, L., Military Space. Great Britain: Brassey’s, 1990. McElroy, J., Space Science and Applications. New York IEEE Press, 1986. Cochran, C., Space Handbook. Alabama: Air University Press, 1985.

1. List and expound upon some of the benefits achieved by using naviga- tion satellites. 2. Assuming a slant range to a GPS satellite of 20,000 km and a speed of light of 3 x lo8 dsec, determine the range inaccuracy associated with a 1 % error in time measurement between transmit and receive of a tim- ing signal. 3. Using the parameters above, determine the time measurement accuracy (percentage and number of seconds) required to achieve a 50-meter slant range measurement capability. PART 3 Spacecraft Systems and Design CHAPTER 8 Spacecraft Systems

The preceding chapters have discussed many of the space science topics involved with the operation of spacecraft and spacecraft applications. The purpose of this chapter is to describe, in more detail, the individual ele- ments which make up a complete space system. Before continuing, we must first define precisely what is meant by a space system. The spacecraft itself is part of a larger system, which includes ground-based telecommu- nication, tracking, data collection, processing, and dissemination elements that are described in the chapter under the heading of Ground Support Sys- tems. Other space-based assets such as communication relay satellites used to link spacecraft data to the ground systems and navigation satellites that provide precise geopositioning information to orbiting satellites can also be an essential part of the larger architecture of our spacecraft system. This constellation of interlocking space-based and ground-based systems are all tied to a system user with a mission goal. Whether the mission applications goal is scientific, military, civil, or commercial, the over-arching space sys- tem, with networked space and ground elements, begins and ends with the user. The user establishes the applications goals and operating protocols at the beginning of the system development process and makes use of the communications, remote sensing, and navigation information produced by the system when it is in place. Our description of this complex end-to-end system begins with a discussion of the spacecraft element by outlining the rudiments of satellite design.

SATELLITE DESIGN

Figure 8-1 identifies the basic elements of a satellite system. In addition to the payload, satellite, and launch vehicle, there are many devices that must be added to the system to complement the payload, such as special downlinks, tape recorders, and cryogenic coolers. The mission and pay- load characteristics also place requirements upon the satellite bus subsys- tems to provide special capabilities in each of the subsystem areas, for Spacecraft Systems 167

PAYLOAD + USER APPLICATION REQUIREMENTS

STORAGE ORBIT DETERMINATION M ISSION/PAY LOAD POWER PECULIAR BUS ,, STATIONKEEPING AUGMENTATIONS DOWNLINK

SATELLITE BUS DESIGN HERITAGE

LAUNCH VEHICLES IMPORTANT COST FACTOR

Figure 8-1. Mission requirements. Many factors affect spacecraft design. example, precise pointing and stationkeeping, and autonomous operation for extended periods of time. This theme is further expanded in the fol- lowing sections that follow the design process illustrated in Figure 8-2 which identifies the salient properties of the mission and payload that drive the design of satellite bus subsystems: 1. Attitude Reference and Control (ARC) 2. Power 3. Thermal Control 4. Orbital Maintenance (station-keeping) 5. Propulsion 6. Data Handling 7. Onboard Computer 8. Telemetry Tracking and Control (TT&C) 9. Structure. In the design process, these subsystems are interdependent, and any design must be iterated so as to take into account interlocking design archi- tecture, power, envelope, and mass interactions. The origins of satellite design are mission and payload characteristics. Mission characteristics establish the environment that the satellite will be subjected to during its lifetime in space as well as the earth observation revisit frequency, range from the satellite to earth, velocity, and other important characteristics, all of which are directly related to the payload 168 Introduction to Space Sciences and Spacecrafi Applications

Mission - Altitude - Inclination - Launchyear - Life span Payload \ - Mass Satellite Subsystems - Power - Attitude reference - Data rate and control - Pointing - Power requirements - Thermal \ - Orbit maintenance - Propulsion - Data handling - On-board computer - TT&C - Structure Figure 8-2. Satellite design flow diagram.

application. Payload attributes such as dynamic mass and envelope dimensions, data collection and transfer rates, power profiles, and point- ing and stabilization requirements also determine the design requirements for the satellite bus subsystems.

Mission

Specification of the six orbital elements is determined by user applica- tions requirements concerning temporal (scene revisit frequency), field of view, solar illumination geometry, and other important considerations. Part of the mission design process includes decisions about the launch vehicle’s role in achieving the final orbit and how much propulsion the satellite design must provide for the satellite to reach its final destination. Specifically, the mission affects the entire design because it establishes the environment and duration for the satellite’s lifetime. The orbit influ- ences the power subsystem by setting the eclipse times and mission life- time that the solar panel output must comply with; the orbital maintenance subsystem by determining the neutral density environment that causes drag and the time duration that the satellite will be exposed to this orbit- modifying force; and the satellite propulsion subsystem by determining the Av necessary to achieve final orbit. Spacecraft Systems 169

Payload

The payload mass and mass dynamics (e.g., scanning antennas and spinning mirrors), pointing accuracy, dynamic envelope dimensions, data and power operating profiles, and other characteristics propagate into many complex aspects of the satellite design. For example, the mass of the payload sets the pace for the calculation of the on-orbit mass of the entire satellite, thus setting the goal for the propul- sion subsystem Av. The ARC subsystem keys on the pointing accuracy required by the most demanding payload element and uses this value to determine the candidate control architectures capable of achieving this accu- racy, then goes on to compute the control system’s mass required to stabilize the satellite in its orbit. The on-off duty cycle and power demands of the pay- load create the profile for the power subsystem that must collect and store energy as the satellite goes in and out of (eclipses) a view of the sun. The thermal subsystem considers eclipse and sun periods and satellite orientation along with the mass and power requirements of the payload to arrive at an estimate of the thermal subsystem mass. The orbital maintenance subsystem uses the mass of the payload to drive the total on-orbit mass and converts this into a satellite design envelope yielding a ballistic cross section. This esti- mation is necessary to compute drag (which varies with the environment pre- dicted during the expected mission lifetime), which in turn determines the fuel necessary to maintain the orbit over the expected mission lifetime. The TT&C and data handling subsystems key on the payload data rate and duty cycle to determine the need for downlink bandwidth and the type and capac- ity of onboard data recorders. The structure subsystem evaluates the needs of the satellite using both static and dynamic mass estimates from the payload and other subsystems to compute the total structural mass. Payload requirements influence and can also be limited by satellite per- formance attributes such as pointing accuracy, structural rigidity, power, and thermal requirements. These and many other important considerations are cross-linked as the designer proceeds from mission and payload spec- ification to a review of each of the satellite bus subsystems.

Launch Vehicle and Site Selection

The launch vehicle may be selected prior to satellite bus subsystem design or after the design has been completed. If the former option is cho- sen, the design proceeds down a path of launch vehicle optimization, and a 170 Introduction to Space Sciences and Spacecraft Applications constant tally of information in the form of mass, envelope, and power “vital statistics” must be tracked indicating the degree of launch vehicle capacity, e.g., envelope and lift, used as the design proceeds through the satellite bus subsystem analysis. Conversely, if the satellite bus subsystem design proceeds without a launch vehicle first being selected (the case for a preselected “common bus” heritage design path), then the resulting satellite design must be matched to an available launch vehicle. Figure 8-3 identifies a few candidate launch vehicles that are available to spacecraft designers. Table 8- 1 provides an example of a typical launch performance database for the Pegasus XL that can be launched from an aircraft platform. Pegasus is a leading launch vehicle candidate for smaller satellites such as SMALLSATs and LIGHTSATs. The following summary discussion of each of the satellite bus subsys- tems serves to introduce the more detailed elements of design associated

ATLAS DELTA SHUTTLE

Figure 8-3. Launch vehicles. While the major types of United States launch vehicles are shown, there are several smaller U.S. launchers available as well as several foreign vehicles. Spacecraft Systems 171

Table 8-1 Typical Launch Performance Database for Pegasus XL Aircraft

Lift Capacity (kg) Inclination (degrees) Altitude (km) 0 28.5 90 98 200 410 395 295 280 400 390 375 275 260 600 370 355 255 245 800 345 335 240 230 1,000 325 310 225 210 1,200 305 290 205 190 1,400 280 265 190 170 1,600 260 245 175 155 1,800 235 225 155 140 2,000 210 205 135 115 Fairing (Lx Diu, m): 1.7, 1.1 Launch Cost ($M):13 Pegasus data from “Commercial Pegasus Launch System Payload User’s Guide” (Rel. 3.00, I Oct I993 ). with each subsystem. The designer has many choices of design architec- ture, including types of stabilization, solar array architectures, combined TT&C and data handling functions, and computer control of subsystem functions. In addition, the design process can select various contemporary and advanced technologies such as the type of solar cell, battery, or propulsion fuel, to name a few. Trade-offs can be tried to improve perfor- mance, to reduce weight and power, or for other considerations until the designer is satisfied with the result. The subsystems are ordered in a logical, interdependent progression, as shown in Figure 8-2. For example, selection of the attitude stabilization technique (spin, 3-axis, gravity gradient) determines the number of solar cells that view the sun at any point in time, which has an important influ- ence on the power subsystem. Selection of the stabilization method will also have an influence on the thermal design. Spin-stabilized satellites evenly expose subsystems to hot and cold space, while 3-axis stabilized satellites can have large thermal gradients caused by specific surfaces exposed to hot or cold conditions for extended periods of time. 172 Introduction to Space Sciences and Spacecraft Applications

Each of the subsystem discussions that follow begin with a review of important mission, payload, and launch vehicle inputs that influence the subsystems calculations.

Attitude Reference and Control (ARC) Subsystem

Depending on the mission, a spacecraft may have varying requirements for pointing accuracies. At some times it may be free to tumble and turn, but at others it may have to pinpoint a discrete location on the earth or deep in space. To do this, the spacecraft must be able to determine its own attitude with respect to some reference, and then to modify this attitude to perform the desired mission. Reference devices include earth horizon sen- sors, sun sensors, star trackers, or even magnetometers which measure the flux lines of the geomagnetic field. Attitude control devices include tiny thrusters, angular momentum storage wheels, gravity-gradient booms, and electromagnetic torque devices. A satellite must determine its attitude with respect to some reference (earth, sun, stars, and/or other satellites) and control its attitude to perform the desired mission. Figure 8-4 provides a simplified block diagram of a typical ARC subsystem. The control logic or computer that accepts input from attitude reference sensors is usually based in space but may include ground-based assets in the control loop, directly or as a backup. The mass and configuration of the ARC subsystem are determined by mission and payload inputs that include the final orbital attitude, payload pointing accuracy, and satellite estimated on-orbit mass. These factors influence the types of disturbances that will most perturb the spacecraft body, a few of which are summarized in Figure 8-5. For example, at geosynchronous altitudes, the effects of solar radiation can impart unbalancing torques on the spacecraft body, as shown in Figure 8-6. It is necessary to select the type of stabilization method that is to be considered. Widely used candidates include: 3-axis, zero momentum 3-axis, bias momentum dual-spin spin gravity gradient These stabilization techniques are listed in descending order according to their precise pointing ability. Spacecrafi Systems 173

EPHEMERIS INFORMATION MAY BE ON ,0 THE GROUND

REFERENCES-HAL [j*(OO '1' CONTROL LOGIC SENSORS ACTUATORS OR ClRCUlTRY 47

L ----_____ 1

OFFSETS CONTROL AND RATES SPACECRAFT TOROUES EOOY

DISTUREANCES

Figure 8-4. Basic attitude reference and control system.

SOLAR RADIATION

MAGNETIC - GRAVITV GRADIENT -

AERODYNAMIC

I 100 1,d, 1o.Ooo 100,m

Figure 8-6. Solar radiation propeller torque. 174 Introduction to Space Sciences and Spacecraft Applications

Figure 8-7 provides a rudimentary comparison between dual spin and momentum bias systems. In spin-stabilized satellites, the larger mass satellite bus spins (approximately once per second in geosynchronous orbits) and a smaller mass (communications antennas, remote sensors) is despun so that it can point in an assigned direction (a point on the earth or other planetary body being observed or communicated with). Conversely, the 3-axis stabilization system controls the spacecraft body with one or more internal spinning masses under the direction of the control comput- er. Figure 8-8 illustrates this even further for a simple body stabilized momentum bias system using a single spinning wheel to control the spacecraft pitch axis while thrusters andor magnetic torquers provide control forces in roll and yaw.

Moderately complex control Moderate accuracy (1 .Onto 0.1’) Can point to a fixed location

II Figure 8-7. Gyroscopic stabilization (dual-spin).

Yaw and roll axes are controlled by thrusters or magnetic torquing Pitch axis is controlled by varying the wheel speed

Figure 8-8. Body-stabilized momentum bias concept. Spacecraft Systems 175

The ARC subsystem takes its most important cue from the pointing accuracy of the single most demanding payload. This determines the avail- able ARC architectures capable of achieving this accuracy: greater than or equal to 5 degrees for gravity gradient, 1 degree for spin, 0.1 degree for three-axis bias momentum or dual-spin, and less than 0.1 degree for three- axis zero-momentum. Having selected an ARC architecture, the designer matches the selected architecture with the minimum components required for implementation as indicated in the example shown in Table 8-2. The mass for the components, required to implement the selected ARC, is arrived at by a Combination of table look-up and calculation. The table look-up mass values are shown in an ARC database such as the one shown in Table 8-3. The contribution to the ARC subsystem mass is simply:

(mass of the component) x (number of components)

As indicated in Table 8-3, the mass for many of the components, such as the reference sensors, is represented by values based on current tech-

Table 8-2 ARC Components Required

~~ ~~~ Stab Method Reaction Momentum MagTor IRU Sun Star Earth Mag 1 3-axis, zero-momentum 4 0 2 03022 2 3-axis, bias-momentum 1 0 2 03022 3 Dual-spin 0 1 2 03022 4 Spin 0 0 2 03022 5 Gravity gradient 0 0 0 02002

Table 8-3 ARC Components Mass (kg)

~ ~~~ S/CMass(kg) Wheels Magtor IRU Earth Momentum Sun Star Mag

~ ~~ 200 2 1 3 2.5 2 0.5 8 1 500 4 1 7 4 4 0.5 8 1 900 6 2 11 5 6 0.5 8 1 1,200 8 3 14 6 8 0.5 8 1 2,000 10 4.5 17 7 10 0.5 8 1 176 Introduction to Space Sciences and Spacecraft Applications nology which must be continually adjusted to reflect the latest available components. The mass of the actuators that provide control torques to counterbalance disturbances or perform maneuvers are a function of the on-orbit mass of the satellite and include the following components: thrusters reaction and momentum wheels magnetic torquers The ARC design begins with a default/minimum value for the number of components needed as a function of the stabilization method, as illustrated in Table 8-2. There are two exceptions: Star sensors are only used in the default result if the pointing accuracy requirement is c 0.1 degree, and magnetome- ters are used only if the spacecraft operating altitude is < 2,000 km. The designer can make decisions based on special knowledge about the application. For example, the designer may add star mappers to a design that does not require them for pointing to provide, after the fact, pointing knowledge that improves data processing and geophysical application. Another option is to increase or decrease the number of components. Increasing the number of components would build in more redundancy, perhaps extending lifetime. The selection of an ARC design architecture impacts the power and thermal subsystems. The most prevalent power subsystems make use of solar cell power generation, and the ARC subsystem determines how the solar cells view the sun. Spin-stabilized or gravity gradient ARC designs do not continuously point the entire solar array at the sun as do the 3-axis stabilized ARC designs. This is taken into account in the power system calculation and determines the number of solar cells, of a particular type, that must be used. Correspondingly, the thermal environment of a 3-axis stabilized ARC design is more demanding than that created by spin-stabi- lized ARC designs, and this will influence the thermal control system mass. As indicated, a spin-stabilized satellite acts like a rotisserie, provid- ing even exposure to hot and cold conditions. A 3-axis stabilized satellite is subject to extreme thermal gradients because particular elements are exposed to hot or cold conditions for extended periods of time.

Power Subsystem These are the components which provide the power used by the space- craft to support itself and perform its mission. Additionally, the hardware Spacecraft Systems 177 used to control the distribution of the power is an important element of this system. Solar, nuclear, and other power-generation sources have been used onboard satellites. Examples of power systems used in spacecraft are described in the following paragraphs. Since solar cells are the most com- mon source of spacecraft power, they will be used in the detailed descrip- tion of the design process.

Solar. There are two types of systems that rely on the sun’s radiations to generate power: solar cells and heat exchangers. Solar cells directly con- vert the sun’s photons (light) to electricity by their reaction to the incom- ing photon flux. Current solar cell technology provides panels capable of generating 140 watts or more of power per square meter of cells. The angle of incidence to the sun is very important and there is a general dete- rioration of the cells of up to 10% per year due to exposure to the space environment. Heat exchangers use the sun’s total radiated energy (the solar constant) to heat a working fluid which then turns a generator to pro- duce electricity. Generated power quantities may be generally higher for heat exchanger systems, but their disadvantage over solar cells is their increased weight and complexity.

Nuclear. There are also two types of nuclear-related power generators which can be used in space: fully critical nuclear reactors and radioisotope thermal generators (RTGs). Nuclear reactors, like solar heat exchangers, use the heat of the nuclear reaction to heat a working fluid which turns a generator to pro- duce electricity. These systems can provide very large quantities of power for missions with high power requirements, such as space-based radars. RTGs also produce heat, but they are noncritical (and inherently safer) and produce lower quantities of electrical power. Both sources are very stable, exhibit long life, can operate far away from solar sources, and can be more compact than solar cells (for maneuvering purposes). Disadvantages are that they are generally very heavy, are radioactive, and are currently in disfavor as a result of adverse public and political reaction to their use.

Chemical. This type of power source includes batteries, with which we are all familiar, and fuel cells in which two elements are recombined with a subsequent release of electrical power such as the oxygen-hydrogen fue’ cells currently used aboard the Space Shuttle. All spacecraft employ bat, teries, at least as storage devices, using the major power-producing sys tem to charge the batteries and subsequently withdrawing power from thc 178 Introduction to Space Sciences and Spacecraji Applications storage device during times when power cannot be generated, such as dur- ing launch, eclipse times, or malfunctions. Fuel cells represent more power availability than batteries as the quantity depends only on the amount of fuel elements carried. The Space Shuttle cells have an interest- ing benefit, as the by-product of the power generation is pure, potable water used by the astronauts for drinking. The disadvantage of both sys- tems is their greater weight versus the amount of power produced com- pared to other methods.

Magnetic. It is also possible to produce power in other ways. One of these could use the process by which a current is produced when a loop of wire is moved through a magnetic field. The earth’s geomagnetic field and a spacecraft’s motion through it represents a possible future space power source. This potential was demonstrated on a Space Shuttle flight by deploying a payload on the end of a 12.5-mile-long tether, however, it has not been used operationally in space as have the other sources of power discussed above.

Solar Cell Power Systems Design

The power subsystem takes the stabilization method as an input cue from the ARC subsystem to determine the area and mass scaling factors that will be used for power subsystem calculations. Solar panel area and mass scaling factor dependency on the selected stabilization method is approximately:

Stabilization Method AredMass Scaling Factor 3-axis, zero momentum 1 3-axis, bias momentum 1 Spin and dual-spin 3.141593 Gravity gradient 4

Using these values as the foundation for the calculation of the design for the power subsystem, the design is a complex, interactive process that proceeds based on a multitude of mission and payload inputs. The electrical power subsystem (EPS) supplies electrical power to the satellite, then stores, regulates, and distributes the power as shown in the block diagram of a typical power subsystem depicted in Figure 8-9. Power subsystems designers calculate the subsystem mass by determining the Spacecraft Systems 179 satellite's end-of-life (EOL) power requirement, working backward to determine the satellite's beginning-of-life (BOL) power requirement. BOL power is higher because solar cells degrade over the mission lifetime (influenced by solar cell type, mission altitude, and inclination), and the solar panels must be large enough to guarantee adequate EOL power. When the BOL power requirements are known, the next step is to size the solar array, batteries, power regulation, control, and distribution systems. Solar cell BOL wattdarea and wattdmass are selected from a database, such as the example provided by Table 8-4 which lists average values for various solar cell types and array architectures. Solar array sizing takes into account the solar cell type sun incidence angle, battery type, eclipse cycle, power duty cycle, stabilization method, and other related factors. Solar eclipse cycle times versus spacecraft circular orbit altitude are illus-

c 7 SERIES SERIES - I REGULATOR I I REGULATOR - DISCHARGE SOUR REGULATOR "zzz ARRAY

(FULL LOAD LOAD OR I PARTIAL) I' SOUR BArnRY ARRAY LOAD

0I OPnONALE= ELEUEMS Figure 8-9. Electrical power subsystem configuration options.

Table 8-4 Solar Cell Design Characteristics

Type Rigid Array Flexible Array Power Mass Power Mass Name (wattdm2) (watts/kg) (wattdm2) (wattskg) 1 SSF Silicon 153 28 153 84 2 ThinSi(APSA) 153 30 153 115 3 GaAsIGE 218 39 218 110 4 CleftGaAs 223 45 223 163 5 GaAsCulnSe2 272 51 212 150 180 Introduction to Space Sciences and Spacecraft Applications trated in Table 8-5. Sun incidence angle is selected by the designer to reflect the maximum incidence angle the solar panel will experience. If the solar panels are sun-tracking, the incidence angle is usually set near zero. The final result is the determination of BOL and EOL power and power subsystem mass.

Table 8-5 S/C Eclipse Times Based on Altitude

Circular Orbit Solar Eclipse Altitude (km) Times (minutes) 200 37.3 250 36.9 300 36.6 350 36.3 400 36.1 450 35.9 500 35.8 600 35.5 700 35.3 800 35.1 900 35 1,000 34.9 1,500 34.8 2,000 35 2,500 35.4 3,000 35.9 3,500 36.4 4,000 36.9 4,500 37.5 5,000 38.1 6,000 39.4 7,000 40.6 8,000 41.8 9,000 43.1 10,000 44.3 15,000 50 20,000 55.2 25,000 60.1 30,000 64.6 35,786 69.4 Spacecraft Systems 181

Several design selections must be made, such as either rigid or flexible arrays and types of solar cells and batteries. Typical battery performance values are shown in Table 8-6. The results of the power subsystem design include EOL and BOL power, determination of solar panel area and mass, number of batteries and their mass, and power control unit, regulator/con- verter, wiring, and total subsystem mass.

Table 8-6 Spacecraft Battery Characteristics

~~~ Specific Energy Density Allowable (SMb) Depth of Battery Type (kglwatt hrs) Discharge (YO) 1 NiCd 0.040 22 2 NiH, 0.025 60 NiCd: Nickel-Cadmium NiH,: Nickel-hydrogen

Power Subsystem Design Calculations. As indicated, mission and pay- load inputs have an important influence on the power design calculation. The final orbital altitude, inclination, and mission lifetime determine the solar degradation rate based on established curves and sets the day/eclipse times that will determine the amount of time the solar cells will be creat- ing power and the amount of time that the payload and supporting systems will rely on battery power. The power required by each element of the payload, the payload duty cycle, and the day/eclipse profile sets the stage for the subsequent system power estimation for the entire satellite. After the baseline power requirements for the satellite are established, the designer makes several important selections that will influence the system mass based on the type of solar array and solar cell desired:

Rigid Array: Derives the mass of the solar array from the solar cell power data base (Table 8-4). Select the Solar Cell Vpe: Derives the size of the solar array based on the solar cell database (Table 8-4). SSF Silicon Single-sided fixed silicon Thin Si Thin silicon GaAs/Ge Gallium arsenide/germanium 182 Introduction to Space Sciences and Spacecraft Applications

Cleft GaAs Gallium arsenide GaAsCuInSe2 Gallium asenide copper indium selenium Select the Battery Qpe: Drives the specific energy density (mass), and depth-of-discharge (DOD) of the batter- ies based on the battery database (Table 8-6). NiCd Nickel-cadmium NiH2 Nickel-h ydrogen

Once the input conditions have been satisfied, the power subsystem designer calculates the basic functional values of the system in the fol- lowing manner:

solar Panel Area (m2> = (BOL Solar Panel Power) x (Area Scaling Factor) (Solar Cell Wattlsq meter)

Solar Panel Mass (kg) = (BOL Solar Panel Power) x (Mass Scaling Factor) (8 - 2) (Solar Cell Kg/sq meter)

The minimum number and mass of the batteries is calculated based on the energy storage needs. The number of batteries may be increased for redundancy.

Min. Number of Batteries = (Total WH Stored) (8 - 3) ([Bus Voltage] x [AWBatt])

Battery Mass (kg) = (Bus Voltage) x (AWBatt) x (8 -4) (Num of Batts) x (Batt Mass to Power Density)

The mass of the power control and conditioning systems is based on a percentage of the power produced.

Power Control Unit Mass (kg) = 0.02 x (Day Power) x 2 (8-5) Spacecraft Systems 183

RegulatordConverters Mass (kg) = 0.025 x (Day Power) x 2 (8-6)

The mass of the wiring is based on a percentage of the overall space- craft mass.

Wiring Mass (kg) = 0.04 x (Estimated S/C Dry Mass) (8-7)

These values define the output of the power subsystem which consti- tutes the largest mass subsystem and one of the more complex elements in the satellite design.

Thermal Subsystem

The thermal control system keeps the temperature of the spacecraft within specified ranges. Too much built-up temperature may affect elec- tronic systems, causing failures, and low temperatures may freeze up movable systems or cause fuel lines to freeze and burst. The control sys- tem must be able to react throughout the expected environmental and operational configurations experienced. Low-earth orbiting spacecraft continuously cycle in and out of the sun’s radiations, absorbing and radi- ating energy in phases. Geostationary satellites may be exposed to (or eclipsed from) solar radiation continuously for weeks and months, and sun-synchronous satellites may be continuously exposed. Internally, a spacecraft may generate large quantities of heat during times of peak operation, and little during periods of dormancy. Thermal control devices fall into two categories: passive and active. Passive devices simply shield, insulate, or change their thermal character- istics depending on the existing temperature of the satellite. The external coating of a spacecraft determines the craft’s absorptivity (how much external energy is absorbed) and emissivity (how much internal thermal energy is radiated into space) characteristics to control temperature. The space shuttle rotates in a maneuver called rotisserie to alternately expose the black underside and the silver-coated open cargo bay doors for the purpose of thermal control. Many spacecraft are wrapped in thermal blan- kets to retain internal heat, and some spacecraft are equipped with pas- sively heat-activated louvers which open or close to expose different external surfaces to radiate or retain internal heat. Active devices usually involve some sort of working fluid to carry heat from one location within a spacecraft to another to either increase or 184 Introduction to Space Sciences and Spacecraft Applications decrease internal or even local temperatures. Refrigeration devices, heat pumps, and heat pipes are examples of active spacecraft thermal control systems. These devices have the ability to more precisely control space- craft temperatures, but their disadvantage is their increased weight and complexity. Special thermal systems, such as cryogenic coolers, are usu- ally accounted for as unique payload support systems. Thermal subsystem mass is usually estimated as a percentage of spacecraft dry mass, typical- ly 2-5% for passive systems and 48%for active systems.

Orbital Maintenance Subsystem

This subsystem design determines the amount of propellant needed to keep the satellite in the desired orbit over the specified mission lifetime. Atmospheric density causes drag and is the primary force that will cause many spacecraft to lose altitude. Atmospheric density is a function of the solar environment, which can vary widely as a function of the expected solar cycle activity, and satellite altitude, as shown in Table 8-7. The final orbit altitude and mission timeline is used to select the appropriate value of atmospheric density. The satellite ballistic cross-section, which is the cross-sectional area of the satellite in the direction of flight, determines the amount of drag. A mission duration Av is calculated to determine the total propellant needed to maintain the orbit. The amount of propellant

Table 8-7 Atmospheric Density

Expected Solar Cycle Activity (kg/m2) Altitude (km) Low Average High 1 00 9.8E-09 9.8E-09 9.8E-09 200 1.95E-10 2.7E-10 3.45E-10 300 8E- 12 2.23E- 11 3.658-11 400 1.04E-12 4.77E-12 8.5E-12 500 1.21E-13 1.05E-12 1.98E-12 600 2.45E- 14 3.12D- 13 6E- 13 700 7.1E-15 9.3E-14 1.8E-13 800 3.4E- 15 3.4E-14 6.4E-14 900 2E-15 1.4E-14 2.5E-14 1,000 1.3E-15 6.7E-15 1.2E-14 Spacecrafr Systems 185 needed is influenced by the specific impulse of the propellant (Isp). The higher the specific impulse, the greater the energy, which results in less propellant needed for on-orbit maintenance. Typical specific impulse val- ues are shown in Table 8-8. Both monopropellants such as N2H4 (hydrazine), and bipropellants such as N2H4/N204(hydrazinehitrogen tetroxide) are used for orbital maintenance. Depending on spacecraft operational altitude, atmospheric drag may be compensated by an electro- static system which produces a small amount of thrust over long periods of time. The details of an orbital maintenance subsystem design depend upon the specifics of the satellite orbit (LEO, SSO, or GEO). For GEO missions, nortWsouth and eastlwest station-keeping requirements must be considered due to perturbations produced by sun and moon gravitational effects as well as nonuniformity of the earth’s gravitational field. The output of the design process is the mass of the orbital maintenance fuel and associated tankage, plumbing, and thrusters. Tankage and plumbing may be shared with the propulsion subsystem; however, in many instances, thruster sizing and placement is unique to the orbital maintenance subsystem.

Propulsion Subsystem

The satellite propulsion subsystem is used to take the spacecraft from a parking orbit to a final orbit. The spacecraft design may choose a “stand- alone” stage for this task. In most cases, modem spacecraft will use a com-

Table 8-8 Thrust System Specific Impulse

Type Propellant ISP Liquid Monopropellant 220 220 Liquid Bipropellant 350 350 Liquid Bipropellant 425 425 Liquid Bipropellant 430 430 Liquid Bipropellant 450 450 Electrostatic Resistojet 400 400 Electrostatic Resistojet 700 700 Electrostatic Arcjet 500 500 Electrostatic Arcjet 1,000 1,Ooo Electrostatic Arciet 1,500 1,500 186 Introduction to Space Sciences and Spacecraft Applications mon or integrated system for orbital maintenance and propulsion and not have a separate perigee or apogee kick motor. Therefore, the specific impulse for both the satellite propulsion subsystem and the orbital mainte- nance subsystem have the same value, and the systems share some common tankage and plumbing. Thruster placement, design, and specific fuel mass may be unique for either orbital maintenance (station-keeping) or orbital translation (propulsion). Since many spacecraft combine these functions, a more detailed description of a composite system is explained in this context.

Propulsion and Orbital Maintenance As a Combined Subsystem

In most cases, onboard propulsion systems must be included to allow the spacecraft to modify its orbit (if necessary), maintain the orbit despite the perturbing forces it may encounter (orbital maintenance), and sometimes, as an adjunct, to control the desired attitude of the spacecraft. To accomplish this, the spacecraft may have a relatively large rocket for orbit modification, smaller rockets for station-keeping, and tiny thrusters for attitude control. Associated with each of these are fuel storage, delivery, and control devices. The propulsion and orbital maintenance subsystems are listed separate- ly based on a design heritage that recognized that many traditional, mod- ular satellite designs had completely separate propulsion and orbital main- tenance systems. For example, some designs used solid-fueled propulsion systems to translate from the parking orbit (achieved by the launch vehi- cle) to the final orbit and a separate liquid-fueled orbital maintenance sys- tem for station-keeping during the life of the satellite. However, many contemporary designs combine the propulsion and orbital maintenance systems using a common fuel and storage system and specific and sepa- rate propulsion and orbital maintenance thruster designs that are unique to the function of orbital translation and orbital maintenance, respectively. The fuel mass is specific for the respective propulsion and orbital mainte- nance categories where they accrue, and will be explained in that way in the following paragraphs. The propulsion and orbital maintenance calculations are both based on the Av or change in velocity (vector) required to perform either an orbit translation or orbital maintenance. Both calculations are based on the equa- tions associated with orbital mechanics that determine the Av required for making a change from one orbit to another. In the case of the propulsion subsystem, the change is associated with going from the parking orbit to the final orbit. In the case of the orbital maintenance system, the designer cal- Spacecraji Systems 187 culates the atmospheric drag on the satellite associated with the precise mis- sion lifetime (calendar years), which invokes the expected spacecraft envi- ronment during this lifetime expressed in terms of atmospheric density as a function of orbit altitude. Atmospheric drag acts in the opposite direction of the orbital velocity vector and removes energy from the orbit. This reduc- tion of energy causes the orbital radius to get smaller, leading to further increases in drag. Acceleration due to drag on a satellite is a function of the atmospheric density, the satellite cross-sectional area in the direction of flight (ballistic cross-section), and the satellite mass (in this case the on- orbit mass). The drag is calculated using the expected environment and the satellite ballistic cross-section. Various shapes and densities exist from the distributed modular design architectures to the dense, highly integrated SMALLSAT designs. This diversity in density and satellite shape results in an individual determination of the ballistic cross-section of a particular satellite. These are all known quantities and are used, in turn, to calculate the Av required to restore the satellite to the prescribed orbit. The remaining quantity required for this calculation is the Isp or specific impulse of the fuel that will be used to perform this maneuver. The specific impulse is a mea- sure of the energy content of the propellant used and how efficiently it is converted into thrust, or how thrust is produced per time rate of change of propellant molecular weight. The Isp of the system to be used by the propul- sion andor orbital maintenance subsystems can be selected by the designer from a database such as the one shown in Table 8-8. The orbit modification propulsion subsystem has to provide significantly more Av than the orbital maintenance subsystem. The propulsion subsystem must provide the thrust to take the spacecraft from the parking orbit to the final orbit. The required performance of the system is expressed as a required Av which is calculated using the principals of orbital mechanics. Usually only the magnitude of the velocity vector is changed (the condition when the parking orbit inclination or plane is the same as the final orbit), but sometimes a plane change is required (the situation for the positioning of most geosynchronous communications satellites). A design must calcu- late the Av required for either one of these situations and, using this and the initial mass of the satellite (m,,) and the Isp of the selected fuel, proceeds to determine the fuel mass (m,) required to perform the maneuver:

where g is simply the gravitational acceleration. 188 Introduction to Space Sciences and Spacecraft Applications

Thruster and tankage mass are calculated using accepted thruster com- ponent and tankage values as a function of the total fuel mass, the number of thrusters required, or other factors. Qpical values are 10% of the required fuel mass. After the mission and payload input has been com- pleted, the designer estimates a spacecraft dry weight (no fuel) and then, using the above method and fuel calculations, makes an estimate of the spacecraft wet (parking orbit) and on-orbit (final orbit) mass estimates. These values are then used to make new estimates for the propulsion and orbital maintenance subsystems. The process is repeated iteratively until an accessible estimate is obtained.

Data Handling Subsystem Most spacecraft are engaged in the collection or transfer of some form of information, even if the specific mission performed by the spacecraft is not involved with communications. The data handling subsystem collects data from the payload and sends the data to the ground. The collected data may be transmitted as it is collected or recorded for later transmission. Data transmission is either direct to a ground station and/or to a ground station via a communication relay satellite, such as the tracking and datu relay satellite (TDRSS). The designer must select the downlink method and frequencies and indicate if onboard recorders are used. A design will calculate the downlink rate (Kbps) and present the downlink options for users to select from. There are several communication bands normally used to send the data to the ground, such as: S-Band (2-4 GHz) Low data rate (could be shared with TT&C) X-Band (8-10 GHz) Payload specific high data rate K-Band (13-15 GHz) Payload specific high data rate After the communication band(s) are selected, the designer calculates the subsystem envelope, mass, and power requirements using a database of transponder, antenna, and recorder information. Conventional designs assume that multiple transponders, antennas, and recorders are always used for redundancy. The result of the data handling subsystem analysis is the determination of subsystem mass and power.

TT&C Subsystem Associated with most spacecraft is the telemetry signal which contains information on the status of the spacecraft and its systems and through which a controlling ground station can modify the spacecraft’s instruc- Spacecraj? Systems 189 tions (commands) for future actions. Modem sophisticated spacecraft are equipped with solid-state control devices on which mission logical instructions can be placed and through which the spacecraft will perform those instructions autonomously. As part of this overall spacecraft coordi- nation, the controller may interrogate its separate systems for their status and perform corrections in cases of malfunction or failure. The telernetrj, trucking, and control (TT&C) subsystem usually has a low data rate (1-2 Kbps) and can do double duty serving as a downlink for low-rate payload data. If the composite payload data rate is low, it can be accommodated by the TT&C subsystem. If the composite data rate is high, a payload specific channel is required in the data handling subsystem. Like the data handling subsystem, a TT&C design calculates the sub- system mass and power requirements from available systems assuming two transponders and omnidirectional antennas for redundancy.

Data Handling and TT&C As a Combined Subsystem

The data handling and telemetry, tracking, and command (TT&C) sub- systems, while separate in the subsystems listing, are associated in much the same way as the propulsion and orbital maintenance subsystems. To conserve weight and space, a design may take advantage of the shared use of telecommunications resources. If an S-band TT&C link can also be shared to collect the payload information being collected, then satellite complexity and cost is reduced. The data handling requirements are based on the payload input values of data rate, duty cycle, and whether or not onboard recording is specified. The data rate will determine whether a low-frequency or high-frequency channel is required for transmission (see the information bandwidth description in Chapter 5) and, hence, whether the low-frequency (S-band) 'IT&C channel can be shared for remote payload data collection. If the data rates warrant, a wider bandwidth channel (X-band or even K-band) will be added as a sepa- rate data handling subsystem. A designer could amend this by taking advan- tage of a low-percentage duty cycle (instrument on time of a few minutes per orbit) and use a recorder playback at a lower rate to return to a shared data handling TT&C design. If the payload input has onboard recording as a requirement, the designer can take this into account in determining the mass of the data handling subsystem. The designer is still responsible for selecting the type of recorder desired (magnetic tape or solid state). Once the component suite is complete, the designer tallies the mass and power requirements for the combined data handling and TT&C subsystem 190 Introduction to Space Sciences and Spacecraft Applications using contemporary component values. The elements include transponder mass, hemispheric and directive parabolic antenna mass, and wiring mass. Typical values for a combined TT&C and data handling subsystem are 2-6% of spacecraft dry mass.

Onboard Computer Subsystem

The onboard computer subsystem asks the designer to specify what function the computer will be used for. The functions include: attitude control navigation control orbital maintenance command processing telemetry processing payload management power management The onboard computer may control functions such as power and data management, recovery from failure, and others. The onboard computer system mass and power are estimated as a function of computer memory storage size (K-words) and throughput (Kbps). The designer indicates which subsystems are communicating with the onboard computer. Based on this selection, a design estimates the storage word size and the through- put and establishes the subsystem mass and power. The designer can also specify whether the satellite is designed with sim- plified or complex ‘‘stand-alone” autonomy. Complex autonomy provides significantly more computer capability because it is associated with systems that are required to make important operational and emergency decisions without benefit of ground analysis or control and may operate for several months without any ground interaction whatsoever. The default design deci- sion is to use full redundancy in this important design area. A designer can change this decision if application goals, i.e., special limited lifetime mis- sions, so warrant. In many designs, mass and power requirements for this subsystem are accounted for with the ‘IT&C and data handling subsystems.

Structure Subsystem

Spacecraft structural considerations must deal with the spacecraft in its many physical configurations. It must be able to support itself under the Spacecraft Systems 191 heavy load of the earth’s gravitational pull, survive the violent process of launch and positioning, be able to deploy systems into an operational con- figuration from the folded, compact configuration of launch, and finally, to maintain spacecraft integrity during station-keeping and attitude changes. The mass of the satellite structure may be estimated as a percentage of the satellite dry mass, typically around 20%.However, the structure mass is also influenced by payload and payload support dynamics, such as scanning antennas and magnetic tape recorder transport, and operating power levels and profiles. An alternate estimation of the mass of the structure subsystem is simi- lar in calculation to the previously described thermal subsystem calcula- tion, with some important exceptions based on uncompensated momen- tum mass imbalance created by scanning payloads. The calculation is iterative beginning with the default values obtained for the mass of each of the spacecraft subsystems, recalculating at the end of the design process. An example calculation might proceed as follows: For each element of the payload the mass multiplier is 0.1 if the pay- load is passive and 0.15 if the payload is active. The payload mass for this calculation is the total mass or the sum of the electronics and sensor mass. If the payload has a scanning antenna that is uncompensated, a compen- sating mass is added to the structure weight. For the ARC and power subsystems structure, the mass multiplier is 0.1 if all of the payloads are passive and 0.15 if any of the payloads are active. The mass multiplier for each of the remaining spacecraft subsystems is 0.1. Each of these contributors are summed to produce the structure sub- systems mass which is allocated to the contributing elements. Results using this type of an approach are consistent with representative designs.

Design Iteration

Since the determination of the mass of a spacecraft subsystem depends upon the mass of each of the other subsystems (for example, ARC mass will depend upon payload requirements as well as the mass of all other subsys- tems), a design must employ an iterative approach for establishing subsys- tem masses. In other words, as each subsystem is designed, all known mass- es are taken into account. Upon completion of the design of all subsystems, an iteration process will be initiated whereby the mass of each subsystem is recalculated, taking into account knowledge of the masses of all other sub- systems. This cycle is repeated until the mass computations stabilize from 192 Introduction to Space Sciences and Spacecraft Applications one iteration to the next. By this process, all subsystems take into account the mass of each of the other subsystems in their mass computations. Table 8-9 provides an overview of the characteristics of some historic and contemporary spacecraft systems. The mission values that drive each design are listed along with some of the top-level characteristics of the satellite bus subsystems. The table provides a perspective about the rela- tive size of subsystem mass values.

GROUND SUPPORT SYSTEMS

Another major component in a space system is the ground support sys- tem. This consists of elements that support the establishment and opera- tion of the spacecraft. Ground systems include the launch, tracking, status monitoring, command, and data collection services required to establish and maintain the spacecraft during its lifetime in orbit. Once the spacecraft is constructed, the support personnel and systems integrate the spacecraft with the launch systems and monitor it before, during, and after the launch, which includes positioning and check-out.

Table 8-9 Comparative Overview of Some Typical Spacecraft Systems

GEOSAT RADAR- LAND- mos- ERS- FO LEWIS CLARK TOPEX SAT SAT6 N 1

Altitude (km) 800 517 475 1,136 800 705 870 800 Inclination (deg) 66 97.4 97.3 98.8 98.8 98.2 98.8 98.8 Launch Year 1996 1996 1996 1992 1995 1993 1992 1992 Life (yrs) 3 5 5 3 3 5 3 3 Satellite Dry Mass (kg) 286 387 278 2,190 3,100 1,750 1,030 2,160 Ballistic Area (m2) 1.2 1.5 1.2 6.5 12.8 5.8 4.1 13 Stabilization 3-AXIS 3-AXIS 3-AXIS 3-AXIS 3-AXIS 3-AXIS 3-AXIS 3-AXIS ARC Sun Sensor YES YES YES YES YES YES YES YES No. of Star sensors 0 2 2 2 0 2 0 2 Mas (kg) 34 38 40 205 156 130 77 156

Power Endof Life(W) 738 740 581 2,140 2,700 1,900 1,500 3,000 Mass (kg) 113 140 104 550 640 380 200 650 Thermal Mass (kg) 7 12 8 54 93 29 46 65

DatflTC Mass (kg) 45 42 45 177 216 325 66 87 Power (W) 66 60 66 0 0 0 0 0 Structure Mass (kg) 30 44 38 480 636 250 217 488 Spacecrafi Systems 193

After the spacecraft is declared operational, the support personnel and systems continue to monitor it for health and operability. Meanwhile, the operations support system has begun to utilize the spacecraft and support this use. This includes ground transmit and receiving stations and data col- lection, reduction, and dissemination systems.

Spacecraft Design and Construction

Depending on the originator of the spacecraft idea, design and construc- tion may occur in many different locations. Some experimental systems may be the product of university-associated laboratories, such as the Johns Hopkins Applied Physics Lab (APL), and the California Institute of Tech- nology’s Jet Propulsion Lab (JPL). The National Aeronautics and Space Administration (NASA) may create some of its own spacecraft at any of its diverse locations, such as the Goddard Space Flight Center (GSFC) or Kennedy Space Center (KSC). Preliminary military space system design may occur at the Air Force’s Space Division or the Navy’s Space Systems Activity (SPAWAR). But, in most cases, operational spacecraft will be designed and built by industrial facilities based on requests for proposals (RFPs) presented by the activity desiring the spacecraft to be built.

Launch Sites and Vehicles

After a spacecraft has been built and tested, it is transported to the launch site to be mated with the launch vehicle for delivery into space. In the United States, the two main launch sites are the eastern test range (ETR) at the Kennedy Space Center in Florida and the western test range (WTR) at Vandenberg Air Force Base in California. Airborne launches from diverse locations have also become available. Table 8-10 lists the total launch vehicle launch weight, lift-off thrust, and maximum payload weight to a typical low altitude orbit for some select launch vehicles. All numbers are approximate. As indicated earlier, launch vehicles are designed simply to deliver their payloads into space and are usually not capable of reaching the final orbits from which many systems operate. Therefore, upper stages may be used to position the spacecraft into (or close to) its operational orbit. As mentioned before, most spacecraft have an associated propulsion system which is used for final positioning and possible repositioning, but the 1% Introduction to Space Sciences and Spacecraft Applications

Table 8-10 United States Launch Vehicles*

Launch Lift-off Low-orbit Weight (Ibs) Thrust (Ibs) Payload (Ibs) Delta 426,000 63 1,100 6,000 292,600 37 8,000 11,200 Titan 15 19,600 2,920,000 34,050 Shuttle 4,500,000 6,425,000 65,000 *Typical perJormances of the mast common US.launch vehicles are given. upper stage is a separate rocket which is usually separated from the space- craft after use. An important point is that if an upper stage is required, it usually sub- tracts from the available weight and space for the payload in the launch vehicle. (Because launch vehicles are multistage rockets, some of the final stages may be capable of delivering the payload into the desired orbit with- out the use of an additional upper stage.) Some common upper stages are the liquid-fueled Centaur (used to send the Viking and Voyager probes away from the earth), and the solid-fueled Payload Assist Module (PAM) and Inertial Upper Stage (IUS), both capable of use from the space shuttle.

Ground Sites

Ground sites consist of the spacecraft control facilities, which monitor and task the system in orbit, and the receiving sites to which the spacecraft may transmit information for dissemination and use. These are many and varied and depend on the operator of the spacecraft. At the receiving sites, infor- mation is collected and put into a form suitable for dissemination to the users. On the ground is also the most important component of a complete space system: the system user. In many cases, the user drives the system design and later operation, as indicated in Table 8- 11 which illustrates the generic personality of communication, navigation, remote sensing, and planetary systems. In many instances, the user only accesses the informa- tion for whatever purpose the spacecraft was designed in the first place. The task force commander at sea only knows that he has instant access to his units, via the satellite communications link, and knows his and others’ positions accurately due to the availability of satellite navigation. Weath- er information in his area of operations is provided by weather processing Spacecraji Systems 195

Table 8-1 1 Satellite Personality of Generic Systems

Communication Navigation Remote Sensing Planetary

~ ______~ Orbit Geosync, , High Earth Low Earth Circular All Types LEO C ir cu 1ar High Eccentric Geosync Power Solar Solar Solar Nuclear, Chem. Solar Attitude Spinning, 3-Axis 3-Axis 3-Axis, Grav Grad 3-Axis Data High Low High + Proc Low Size Moderate to Large Small 500 kg Moderate to Large Small to Moderate 1,000-5,000 kg 1,000-5,000 kg 500-5,000 kg Mission Data Relay X-Mit Position Collect Info various Lifetime 8-12 Years 8- 12 Years 3-5 Years Years cost $50-80M $5-30M $50-500M $0.1-5B Complexity Moderate Low High New Horizons Lead Time 2 Years 2 Years 4 Years 5 Years and dissemination ground sites that allow him to best utilize his airborne, surface, and subsurface assets. Every time you pick up the phone and make a long-distance call or watch cable TV at home, you become a user of spacecraft systems.

SPACE SYSTEM EVOLUTION

The space system design process has evolved over the nearly four decades in which it has been conducted. Table 8-12 provides a historical perspective of the evolution of spacecraft systems over this period. Early space systems were small (due to the existing launch vehicle constraints) single-purpose systems. As launch capabilities increased, so did the size and complexity of spacecraft designs, and billion-dollar, multimission spacecraft have been built and launched. Lessons were learned with these systems when small failures of single systems either decreased the effec- tiveness of the spacecraft or rendered it totally useless. Modem spacecraft design uses both multiple-use, modular construction spacecraft designs with interchangeable systems using more common (and inexpensive) off-the-shelf equipment and the recent resurgence of smaller, single-purpose, highly integrated spacecraft that bear the mantle of SMALLSATs and LIGHTSATs. More and more spacecraft are being 196 Introduction to Space Sciences and Spacecrafi Applications

Table 8-12 Evolution of Satellite Systems

196OSnOs 197Od80s 1990s/2000s

SPACE Small Large Very large, Small, single use SEGMENT Integrated Modular multiple-use design Expendable Modular, expandable Expendable Long life with Short life maintenance and modification

MISSION Navigation, Navigation, Many missions using Single missions communication, communications, one platform (e.g., passive remote active remote space shuttle, space sensing sensing station) Single missions Integrated missions

CONTROL Ground-based Ground- and Autonomous and SEGMENT Low data rate space-based space-based Frequent blind High data rate Onboard processing, periods Selective ground evaluation, Spotty tracking network and decision-making ground stations space relay Continuous coverage Nearly Excellent continuous continuous tracking coverage Good tracking, varying with orbit

designed with the capability for in-space repair or retrieval for refurbish- ment and reuse. The future will see major space systems actually con- structed and tested in space before deployment and operation.

REFERENCES/ADDITIONAL READING

Pisacane, V. and Moore, R., Fundamentals of Space Systems. New York: Oxford University Press, 1994. Wertz, J. and Larson, W., Space Mission Analysis and Design. Boston: Kluwer Academic Publishers, 1991. Griffin, M. and French, J., Space Vehicle Design. Washington, D.C.: American Institute of Aeronautics and Astronautics, 1991. Spacecraft Systems 197

Kaplan, M., Modern Spacecraft Dynamics and Control. New York: John Wiley & Sons, 1976.

1. A geostationary communications satellite requires 1,OOO watts of power to serve its payload and spacecraft bus subsystems. Calculate the number of SSF Silicon solar cells required for a 10-year design life for a spin- stabilized spacecraft. Compare this result to a 3-axis stabilized design. 2. Estimate the power subsystem mass for each of the spacecraft designs in Exercise 1 above. 3. The on-orbit mass of a satellite is 1,000 kg. The satellite begins life in a 66" inclination, 500 km altitude parking orbit achieved by the launch vehicle. Calculate the mass of propellant required to place the satellite into its final orbit of 99" inclination and 1,000 km altitude using bipro- pellant 350. 4. A satellite with a ballistic cross-section of 4 m2 is in an 800 km altitude, 108" inclination orbit where it is expected to perform for 5 years. Cal- culate the fuel mass required using monopropellant 220. Compare this result to the case in which the satellite's orbital altitude is 300 km. CHAPTER 9 Spacecraft Design

In the complex world of aerospace vehicle design, it is no longer possi- ble for a single designer to individually create a vehicle which can opti- mally and capably perform the many functions that modem expectations place on such systems. The word “system” is not used lightly here, as more and more often, modem-day systems are complex assemblages of many different parts. With this in mind, the engineering definition of a “system” may be introduced: A system is a set of elements organized to perfom a set of designated functions in order to achieve specified results. This definition itself intro- duces additional terms which, in turn, must also be defined: An element of a system is an identifiable component of the whole that per- forms some identifiable function. Each element performs particular designut- edfinctiuns which represent the defined actions, operations, or use of the ele- ment. These well-defined actions assist the system in achieving its specified results which are simply the stated objectives or purpose of the system. Thus defined, it is easy to see that an element may be a system in its own right. More than just individual pieces of a particular mechanism, system elements may also include personnel, facilities, information, and many other factors that are not necessarily physical parts of the mecha- nism under design. In viewing the system as a set of many diverse components, it can eas- ily be seen that a single designer, though competent in one or even a few fields, may be lacking in some knowledge that may not allow him or her to compose the optimum method for achieving the particular objective. Designs of large-scale, complex modern-day systems is a multidiscipli- nary task calling on the expertise of many toward the one objective.

THE SYSTEMS APPROACH

In order to coordinate the efforts of those contributing to the task, an organized method must be followed to ensure success. The systems Spacecraft Design 199 approach described herein, or something similar, is the methodology used to approach the problems encountered by NASA and the aerospace indus- tries in meeting the needs of the space programs, by AT&T in meeting the communications needs of millions, and even by urban planners consider- ing public transportation, pollution, health care, and other needs. From the above discussion, it should be clear that the prerequisite for considering the design of a space-based system is that an appropriate mis- sion need exists. A suitable commercial, military, civil, or scientific need must be well established to justify the time, effort, and sizable expense required to warrant the development of a space-based system. The need is often defined by individuals or groups of potential systems users scattered all over the world. Sources of demands for satellites originate from mili- tary or civil government agencies, scientific organizations, and commer- cial service industries. Some space systems may be sovereign in nature, providing information and services to a select but important sector of users. In some cases, international organizations, both scientific and com- mercial, form shared programs with many entities providing funds and systems elements. Advisory committees, such as the National Academy of Sciences and the National Space Council, and universities provide guidance and play important roles in defining needs for manned and unmanned satellites. Whatever the source of the initial concept for the mission or the selection of the specific implementation approach of the system, further development usually follows a fairly common, relatively standardized and structured approach, referred to as the systems approach. As shown in Figure 9-1, the systems approach is a two-dimensional method which has the design process along one axis and the systems life cycle along the other. When placed together, the result is a matrix which delineates a methodical system design process. Each step in this process is explained in the following sections.

System life Cycle

The system life cycle begins with formalization of the mission need and terminates when the system is retired or fails. The life cycle may be orig- inated as a result of a new need, new technology, or as an iteration of a previous system whose life cycle is nearing completion due to obsoles- cence. As indicated in Figure 9-1, the system life cycle has three distinct periods: planning, acquisition, and use. Each of these periods is divided 200 Introduction to Space Scieplces and Spacecraft Applications

Figure 9-1. Systems approach. A structured approach is used in the design of complex systems such as spacecraft. into a number of phases through which the designer conducts the distinct steps of the systems approach. The initial period in the life cycle is the planning period. During this period, the need for the system is established, system concepts are formu- lated, and their feasibility and worth are established. The output of the planning period is an identification of a system concept and a set of sys- tem requirements. The acquisition period includes those steps necessary to design, test, and evaluate the system, and to produce and install it, The use period consists of those activities and resources required to operate, support and maintain the system. The phases associated with the planning period consist of the concepru- al study phase and the mission analysis phase. The acquisition period starts with the system definition phase and ends with the system development phase which includes launch and initiation of operation. Within these two periods, the activities associated with systems engineering are defined. It is during these periods that maximum benefit may be realized in following the design process of the systems approach. As shown in Figure 9- 1, these are mainly applicable to the planning phases and the phases concerned Spacecraft Design 201 with design of the system prior to production. The planning, acquisition, and use periods represent time-based periods in a system’s life cycle.

System Design Process

While the system life cycle identifies the different periods and phases in the design of a complex system, the design process defines the funda- mental sequence of activities used for making design decisions (trade- offs) during each phase. It is an iterative (repeated) method that results in a successively more detailed and optimized solution to the problem at hand. The process is continued repeatedly until either an acceptable solu- tion is reached or a decision is made to regress to a previous design phase (or even back to square one!). The process is broken into a number of stages which are also indicated in Figure 9-1. These stages are discussed in more detail in the following section, after which the phases in the sys- tem life cycle will be presented in detail.

Design Process Stages

Gather Information. In the first stage, the designer (or group) gathers all the information that may be pertinent to the problem. This includes the inputs to the particular design phase (defined later), as well as overall sys- tem constraints and purposes, which may be useful in establishing criteria to which the inputs will be compared and optimized.

Value Model. The next stage uses the information gathered in the first stage to formulate a function (model) that indicates the desired perfor- mance of the system or element with respect to all pertinent considera- tions (such as size, cost, and effectiveness). It is to this model that the evaluation functions of all the possible approaches will be compared. In general there are three important areas to consider: Effectiveness: ability to perform the stated objectives. This mainly deals with the physical capabilities of the solution. Cost: resources required by the solution. Time: when the solution is required or how long it may take to pro- duce the particular solution. Many other considerations may be involved such as risk, politics, and others directly associated with the particular need being addressed. These 202 Intmduction to Space Sciences and Spacecraft Applications considerations are not to be isolated from one another, and each may con- tain a number of factors within itself. Each factor must be identified and the relative importance of each specified in the evaluation function by use of weighting factors. An example may be given for the construction of a road for which we would like to minimize the time between destinations and maximize the efficiency (traffic control), but, of course, minimize the cost. The parameter may look like:

ROAD(tirne x w1, efficiency x w2,cost x w3) (9-1 ) where time, efficiency, and cost represent the important variables of the parameter and wlr w2, and w3 represent the weighting factors assigned to each variable. These variables must all be considered, but some may be more important than others and will be given higher weighting factors in the actual parameter statement indicating their priority. Some of these vari- ables may also conflict with the desires of some of the others. For instance, minimizing the travel time on our road by cutting a hole through a moun- tain may not fit into the cost restraints with which we also must ded. A considerable amount of effort must be put toward this stage of the design process since it directly affects the outcome of evaluation of pos- sible solutions to the problem.

Alternative Solutions. Having decided on the desired performance dur- ing the last stage, the designer may now use his knowledge and analytical skills to synthesize solutions to the problem. Formulation of solutions prior to the last stage should be avoided to prevent biased ideas. This is the fun part of systems engineering where the designer may dream up exotic solutions to the stated objectives and have them compared to the more conventional solutions that will also be brought forward. (Who knows if the exotic solution won’t out-perform the conventional solutions when the evaluation stage is reached!)

Test or Analysis. Before a good evaluation can be performed, as much information as possible on each solution must be obtained. In the test or analysis stage, each solution is tested (using models or simulations), if possible. Alternately, its performance in each of the value model factors may be estimated. This will result in a performance that can be compared to the desired performance model formulated earlier.

Evaluation. The results of the test or analysis stage for each solution are compared to the desired performance function as well as to the results for Spacecraft Design 203 each of the solutions considered. It must be realized that none of the pos- sible solutions may meet all the criteria specified in the desired perfor- mance function completely or to the level desired, so comparison of the solutions obtained may be beneficial.

DecisiodOptimization. At this point, the designer must make a decision. This process is simplified in the systems approach by the fact that only four courses of action are possible: One (or more) of the solutions meets an acceptable level of perfor- mance. The solution which gives the best level is usually the one that will be passed on as the solution to the next phase of the design process, though more than one may be passed on as alternative solu- tions as well. None of the solutions represents an acceptable solution, but one of the alternatives exhibits promise if reevaluated or refined within the same phase of the design. This may involve refining the solution itself in terms of its configuration or analysis, or may require redefinition of the evaluation function itself. This process of feeding back to a stage within the same step of the design process is known as optimization. None of the solutions are acceptable and a return to an earlier stage is required. There is no solution and the program should be dropped.

Communication. The last stage of the design process involves communi- cation of the results. This must be in a form suitable for use by the follow- on activities and may involve reports, drawings, specifications, or models. This represents the end of the design process for a particular step, which may be simply the end of a single phase of the system’s life cycle, or may represent the final step in the overall system engineering design. In any case, the design process is carried out over and over until the program is terminated or the system reaches the end of its life cycle where, if a need still exists, the entire process may be started again.

System Life Cycle Phases

Figure 9-2 shows the different phases of the life cycle for a typical space system development. Each phase of the system life cycle has a def- inite input which, when subjected to the steps of the systems design process, leads to the outputs for the phase and ultimately for the period as 204 Introduction to Space Sciences and Spacecraft Applications well. Each phase is associated with the technical and programmatic (cost, risk, and schedule considerations) content and level of detail expected at that stage of mission development. As seen in the figure, a number of for- mal, documented reviews are used to ensure that the development is pro- ceeding as expected and that all factors are such that the development can proceed to the next level. Considering the expense associated with such missions, and the usual difficulty of retrieving the satellite once placed into orbit, these factors include budgetary and political as well as quality assurance and technical considerations. The following sections will describe the level of detail expected, usual activities, and the reviews and/or decision gates associated with each phase in the unfolding devel- opment of a typical satellite mission.

Conceptual Study Phase (Pre-Phase A). This initial conceptual study phase is usually initiated by the intended users and mission advocates with the assistance of a spacecraft development organization (Le., a NASA center, military organization, or spacecraft company). The phase begins by formulating a clear statement of the mission needs and objectives. Associated with the mission needs statement is a justification for the mis- sion describing the scientific benefits that will be achieved, military goals accomplished, or projected profit to be made. Using descriptions of the

Pre-Phase A Phase A PhaseB Phasem Phase E

Concephlal Mission operations Study phase Analysis phase Definition Phase Design/Dsvelopment Phase phase I Initial Objectives Initial Requirements Design To Detailed Design I Fabrication SpaCW3fl Requirements Operations fustificarion/Benefirs Payload Test Models I lnlegration DefinitiodSelection Reliminary Design Payload implementation Conceptc * Instruments ' Test Operations Concept Analysis/ * Spacecraft I ROM Cost Estimate Selection * Launch Vehicle 1 Launch & Operations Science Systems Engineering * Ground Systems I Preparation Planning Announcements of - Operations 3pportunity Updated Cost & Schedule I Data Collectioi Eitimatcs Scope of Work I Data Reductior Reasonable Schedule I I End-of-Life Expected Costs I Requests for Proposals I

Figure 9-2. System life cycle. The design of a spacecraft passes through many phases of evolution and review before becoming reality. Spacecraft Design 205 mission and payload, a number of different concepts may be formulated that can achieve the mission objectives, and these concepts will be expanded to determine the feasibility of conducting the mission. This con- cept development will include ground system and operations as well as launch vehicle and spacecraft descriptions, but the level of detail is usual- ly in the form of estimates based on contemporary systems and technolo- gies. After the concept is formulated, a rough order ofmagnitude (ROM) cost for the mission is attained using heritage cost models and known costs for subsystems and services. Exceptions to this are unique, one-of- a-kind designs or pioneering first efforts. The results of this phase are reviewed at a “local” level-within a tech- nical center, military or industrial laboratory, university, or similar orga- nization-to determine if the concept warrants further consideration out- side the group. For example, within NASA, if the concept looks promising, an announcement ofopportunity (AO) may be issued which is the first step in defining and selecting the instruments/experiments that will fly on the mission if it is fully developed and launched. If concept study results prove to be promising, the evolving program moves on to the next step in Phase A, a more detailed mission evaluation.

Mission Analysis Phase (Phase A). The purpose of the mission analysis phase is to translate the broad mission concepts and objectives into a fea- sible preliminary system design. This is a refinement and expansion of the conceptual study phase made with the intention of providing a concise, clear overview of the proposed system. Implementation planning is done in cooperation with discipline engineering groups (attitude control, power, thermal, and other spacecraft subsystems areas defined in the pre- ceding chapter), payload specialists (e.g., remote sensors and communi- cations transponders), and ground and launch services organizations. Financial analysts utilize parametric cost models to derive top-level resource estimates for the overall mission and its component elements. Key activities during this phase are system and subsystem trade studies, analyses of performance requirements, identification of advanced technol- ogyAong lead items, risk assessments, and end-to-end system life-cycle costs as a trade parameter. In addition, considerable attention is given to schedule options and selection and evaluation of operational concepts. Specific tasks to support this phase may include technical feasibility and risk assessments associated with cost estimate modeling. Special attention is given to interface identification, definition, and analysis 206 Introduction to Space Sciences and Spacecraft Applications

including spacecrafdground data communications interfaces, satellite tracking system interfaces, and payloadspacecraft bus interfaces. The definition and evaluation of alternative strategies for the execution and management of the mission is also an important consideration. The typical satellite Phase A study has a duration of approximately nine months to two years. Study costs are usually one to two percent of total mission costs.

Definition Phase (Phase B). The purpose of this phase is the refinement of the mission and system architecture and the system design created dur- ing the mission analysis phase. It is designed to convert the conceptual system design, which may still include some open trade-off and develop- ment options, into a final design and baseline for entering the execution phase. Functional, operational, and performance requirements are refined. Established interface requirements and specifications are allocated down to the subsystem or major component level. Firm cost estimates and schedules are prepared to assure a smooth transition to the subsequent execution phase C/D. Once a detailed baseline configuration that satisfies all the mission and program requirements has been established, the defin- ition phase design, cost, and schedule are submitted to the program spon- sor for approval. Key activities during this phase are revalidation of mission require- ments and system operational concepts. This phase also marks the con- version of mission requirements to much more specific project require- ments. Final risk assessments are made along with more narrowly drawn system and subsystem studies and trades. Schedules and life-cycle costs are finalized to support the budget acquisition process. This phase covers the full range of technical, management, resource, facility, and procurement assumptions. Phase B results in the full docu- mentation of the entire proposed flight and ground support system. Alter- nate designs are analyzed in detail to allow the choice of one optimum flight and ground system approach considering technical performance, cost, risks, and schedules. The baseline design is totally defined as well as all major interfaces and system, subsystem, and component specifications. Specific tasks conducted during this phase include telemetry assess- ments, analysis of payload hardware, launch vehicle interface analyses, and in-depth evaluation and categorization of the expected space environ- ment. Components and subsystems are evaluated, as are production plans, facility requirements, and contingency workaround or offset plans. Spacecraft Design 207

This phase has a duration of from one to two years. The costs associat- ed with this study are usually four to eight percent of total mission costs.

DesigdDevelopment Phase (Phase 0).In many instances, these phas- es are combined into a single phase, since by now the commitment to build the system may already have been made. The purpose of the designldevelopment phase is the implementation of the Phase B Plan. The final design, production development, and fabrication of hardware and software needed to translate the mission into a reality begins. The phase often includes launch, deployment, and postlaunch on-station system checkouts and follow-on operations. Depending on the scope of the deployed system (Le., multiple satellite weather or navigation system developments versus one-of-a-kind single satellite programs) this phase can have an extremely variable duration, lasting from two years to more than a decade. The costs associated with this phase can also vary consid- erably-from tens of millions of dollars for SMALLSAT or LIGHTSAT missions to billions of dollars for defense or planetary exploration satel- lite programs. Key activity categories during this phase are linked to the scheduled sequence of events, with each item listed below applicable to all the mis- sion elements (space and ground systems and subsystems; payload ele- ments including spacecraft/payload/ground system interfaces and links): 1. Design and development. 2. Fabrication. 3. Assembly. 4. Integration. 5. Test and evaluation (ambient and space environment exposures). 6. Transfer to the launch site followed by mating and prelaunch checkout. 7. Launch and launch sequence monitoring. 8. On-orbit checkout. 9. Postlaunch operations, which can include retrieval and on-orbit change or repair. Perhaps the most important milestones in the Phase C/Dprocess, shown in Figure 9-2, are the preliminary design review and the critical design review (PDR and CDR). These signify the release of the design for final production review (in the PDR), and the essential freeze of the design for fabrication and assembly (in the CDR). 208 Introduction to Space Sciences and Spacecraft Applications

Specific activities needed to support the sequence of events listed above include: Total quality management for reduction of mission problems (e.g., performance assurance, safety, reliability, maintainability, and quality assurance analyses and assessments). Assessment of specification packages to conform to all applicable standards. Evaluation of technical, progress, and feasibility reports. Liaisodcoordination services to support project activities. Interface documentation and interface engineering. Mission plan. Operations and training manuals. Evaluation of configuration changes and changes to proposed designs or procedures. Monitoring fabrication of all hardware. Monitoring telemetry and command signal margins. Evaluation of logistical support for flight and ground support equipment. Review of functional test plans and procedures. Test and evaluation of the spacecraft, payload, and ground system. Flight and ground software verification and validation. Oversight of launch and operations support. Control of costs, schedules, and technical requirements. Plans for monitoring and controlling the allocations and error budget associated with power, thermal, attitude, timing, and mass properties, including launch margins. Readiness of systems for launch. Monitoring flight operations. Spacecraft/ground system command and data handling coordinated with the mission plan. Assessment of spacecraft/payload problems. The outputs for this phase comprise all aspects of the space and ground segments of the mission and their operation. Prior to proceeding to the operational phase, the launch must have been successful, the spacecraft should be deployed and in the correct orbit, the spacecraft and ground systems communications should be in normal working order, and the spacecraft should be operating successfully in con- cert with the tracking stations, telemetry and command, and data collec- tion facilities. Spacecraft Design 209

Operations Phase (Phase E). The purpose of the operations phase is to safely and productively conduct the routine operation of the satellite, including data acquisition, processing, and distribution, until or unless a problem develops or the mission is terminated. The conduct of this phase varies considerably depending on whether the mission is a military, civil, scientific, or commercial endeavor. Key activities during this phase are command and control of the space- craft and instruments and data transmission, capture, processing, archiv- ing, and distribution. Special mission operations may be authorized and conducted. Orbital changes may be required. Spacecraft or instrument components may be retrieved, repaired, or changed out in orbit. During this phase, routine items include: Payload performance reviews with users. Flight assessments of hardware, software, orbit behavior, and mission objectives. Monitoring the acquisition, processing, dissemination, and archiving of data. Compilation and production of a “lessons learned” document for the mission.

The following is a list of factors that must be considered in the design of systems which are to perform in the environment of space. The list is in no way complete or exhaustive but represents a minimum of topics which must be considered in such systems. These factors were presented in detail in Chapter 8 along with their associated design considerations. A. Space Component 1. Payload (mission-specific) systems 2. Support systems a. power systems b. thermal control c. propulsion d. attitude reference and control e. data f. spacecraft monitoring and control g. structure 210 Introduction to Space Sciences and Spacecraft Applications

B. Ground Support Components 1. Spacecraft design and construction 2. Launch sites a. launch vehicles b. upper stages 3. Ground Sites C. User Components

REFERENCES/ADDITIONAL READING

Wertz, J. and Larson, W., Space Mission Analysis and Design. Boston: Kluwer Academic Publishers, 199 1. Griffin, M. and French, J., Space Vehicle Design. Washington, D.C.: American Institute of Aeronautics and Astronautics, 1991. The Flight Projects Directorate Project Management Handbook, NASA, Goddard Space Flight Center, 1994.

EXERCISES

1. Choose one of the common uses of systems in space (communications, navigation, remote sensing) and complete the following: a. Define the need for such a system and justify its performance from a space-based solution versus a ground-based system. b. Define the top-level characteristics (requirements) for the system. This is a listing of what must be accomplished to meet the stated need, but does not assign functions to any particular segment of the system (space segment, ground segment, etc.). c. Create a value model using these characteristics. Assign and provide rationale for weighting factors for each characteristic. d. For each of the space system components listed at the end of this chap- ter, take a first cut on assignment of functions and component charac- teristics. This will involve defining the basic mission scenario, includ- ing launch site and vehicle, ground systems and operation, as well as defining the types of spacecraft subsystems appropriate for the mis- sion. Assumptions may be made for the payload (size, mass, power, data, etc.). e. The next step would be to perform an initial sizing of the system and subsystems. Good luck! APPENDIX A Manned Spaceflight Summary

The following table summarizes United States and Soviet manned space launches up to the first launch of the Space Shuttle.

Designation Launch Recovery Astronauts Notes

Vostok I 12 Apr 61 12 Apr 61 Gagarin 1 orbit, 1.8 hr. Mercury- 5 May 61 5 May 61 Shepard Suborbital, 15 min. Redstone 3 Mercury- 21 Jul 61 21 Jul61 Grissom 2nd U.S.suborbital Redstone 4 2 6Aug61 7Aug61 Titov 17 orbits, 25 hrs. Mercury- 20 Feb 62 20 Feb 62 Glenn 3 orbits, 4.9 hrs. Atlas 6 Mercury- 24 May 62 24 May 62 Carpenter 3 orbits, 4.9 hrs. Atlas 7 Vostok 3 11 Aug 62 15 Aug 62 Nikolayev 64 orbits, 94.4 hrs. Vostok 4 12 Aug 62 15 Aug 62 Popovich Came within 5 km of Vostok 3. Mercury- 3 Oct 62 3 Oct 62 Schirra 6 orbits, 9.2 hrs. Atlas 8 Mercury- 15 May 63 16 May 63 Cooper 22 orbits, 34.3 hrs. Atlas 9 14 Jun 63 19 Jun 63 Bykovsky 81 orbits, 119.1 hrs. 16 Jun 63 19 Jun 63 Tereshkova Came within 5 km of Vostok 5. First woman in space. 1 12Oct64 13Oct64 Komarov 16 orbits. Feokistov First multiple crew. Yegorov Voskhod 2 18 Mar 65 19 Mar 65 Leonov First EVA Belyayev (Leonov, 20 min.)

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21 1 212 Introduction to Space Sciences and Spacecraft Applications

Designation Launch Recovery Astronauts Notes

Gemini 3 23 Mar 65 23 Mar 65 Grissom Orbital maneuvers. Young 3 Jun 65 7 Jun 65 White 21 An. EVA McDivitt (White). 21 Aug 65 29 Aug 65 Cooper 128 orbits, Conrad 190.9 hrs. Gemini 7 4 Dec 65 18 Dec 65 Boman 220 orbits, Lovell 330.6 hrs. Gemini 6 15 Dec 65 16 Dec 65 schirra Rendezvous (RVZ) within Stafford 1 ft of Gemini 7. Gemini 8 16 Mar 66 16 Mar 66 Armstrong Docked with Agena but Scott suffered control failure. Gemini 9 3 Jun 66 6 Jun 66 Stafford Fairing prevented docking. Cernan Conducted RVZ and EVA. Gemini 10 18 Jul66 21 Jul66 Young Docked and used Agena to raise Collins orbit altitude. Gemini I1 12 Sep 66 15 Sep 66 Conrad Docked with Agena Gordon within 1st orbit. Gemini 12 11 Nov 66 15 Nov 66 Love11 63 orbits, 94.6 hrs. Aldrin Multiple EVAs. Apollo I Grissom Astronauts killed in fire on White launch pad during tests Chaffee 27 Jan 67. soyuz I 23 Apr 67 24 Apr 67 Komarov Cosmonaut killed during recovery. Apollo 7 11 Oct 68 22 Oct 68 Schirra 163 orbits, Cunningham 260.2 hrs. Eisele 28 Oct 68 30 Oct 68 Beregoyov RVZ to 198 m (650 ft) of unmanned . 21 Dec 68 27 Dec 68 Borman 10 lunar orbits. Lovell Anders soyuz 4 14 Jan 69 17 Jan 69 Shatalov Docked with Soyuz 5. soyuz 5 15 Jan 69 18 Jan 69 Vol ynov EVA transfed 2 cosmonauts Khrunov to Soyuz 4 for landing. Yeliseyev Volynov recovered alone. Apollo 9 3 Mar69 13 Mar 69 McDivitt Tested lunar module in Scott earth orbit. Schweickart Apollo 10 18 May 69 26 May 69 Young Lunar landing rehearsal piloted Stafford LM to 15 km (9.3 mi) of Cernan lunar surface. Manned Spaceflight Summary 213

Designation Launch Recovery Astronauts Notes

Apollo 11 16 Jul69 24 Jul69 Armstrong Sea of Tranquility, 21.6 hrs. Aldrin on lunar surface. 2.5 hr. Collins lunar EVA. Soyuz 6 11 Oct 69 16 Oct 69 Shonin 3 simultaneous flights (Soyuz Kubasov 6.7, and 8). Conducted welding tests. soyuz 7 12 Oct 69 17 Oct 69 Filipchenko Navigatiodphoto Volkov experiments. Gorbatko Soyuz 8 13 Oct 69 18 Oct 69 Shatalov RVZ with Soyuz 7. Yeliseyev Comm.lphoto experiments. Apollo 12 14 Nov 69 24 Nov 69 Conrad Ocean of Storms, 3 1.6 hrs. Bean on lunar surface. 2 EVAs. Gordon Recovered Surveyor parts. Apollo 13 11 Apr 70 17 Apr 70 Love11 O2 tank rupture crippled Swigert service module. Haise Crew in LM circled moon to return. soyuz 9 2 Jun 70 19 Jun 70 Nikovayev 17 days, 16 hrs., 59 min. Sevastianov Apollo 14 31 Jan 71 9 Feb 71 Shepard Fra Mauro area, Mitchell 2 EVAs. 9 hrs. Roosa Salyut I 19 Apr 71 11 Oct71 Prototype space station. soyuz 10 23 Apr 71 24 Apr 71 shatalov Docked with Salyut I but did Yeliseyev not enter. Rukavishnikov soyuz I I 6 Jun 71 29 Jun 71 Dobrovolsky Manned Salyut I (22 days). Volkov Crew perished during recovery. F’atsayev Apollo 1.5 26 Jul71 7 Aug 71 Scott Hadley-Apennines area. Irwin Used lunar roving vehicle. Worden Left satellite in lunar orbit. Apollo 16 16 Apr 72 27 Apr 72 Young Descartes area. Duke EVA on return from Mattingly moon to retrieve film cannister. Apollo 17 7 Dec 72 19 Dec 72 Cernan Taurus-Littrow area. Last lunar Schmitt mission. Schmitt first scientist/ Evans astronaut. Salyut 2 3 Apr 73 28 May 73 Space station (failure). Skylab I 14 May 73 11 Jul79 US. space station.

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Designation Launch Recovery Astronauts Notes

Skylab 2 25 May 73 22 Jun 73 Conrad 28 days. Conducted unsched- Kenvin uled EVAs to repair Skylab Weitz damage during launch. Skylab 3 28 Jul73 25 Sep 73 Bean 59 days. More Skylab repairs. Garriott Extensive earth and sun Lousma observations. soyuz I2 27 Sep 73 29 Sep 73 Lazarev Test of Soyuz redesign. Makarov Skylab 4 16 Nov 73 8 Feb 74 CZUT 84 days. Longest Gibson U.S. space mission to date. Pogue Soyuz 13 18 Dec 73 26 Dec 73 Klimuk Astronomical experiments. Lebedev Salyut 3 24 Jun 74 24 Jan 75 Space station. Soyuz 14 3 Jul74 19 Jul74 Popovich Manned Salyut 3. Artyukhin Soyuz 15 26 Aug 74 28 Aug 74 Sarafanov Rendezvous but did not man Demin Salyut 3. Soyuz 16 2 Dec 74 8 Dec 74 Filipchenko Test of ASPspacecraft. Rukavishnikov Salyut 4 26 Dec 74 2 Feb 77 Space station. soyuz I7 10 Jan 75 9 Feb 75 Gubarev Manned Salyut 4. Grechko SOYUZ18-A? 5 Apr 75 5 Apr 75 LaZWV Booster failure aborted mission Makarov after launch. Soyuz I8 24 May 75 26 Jul75 Klimuk Manned Salyut 4. Sevastyanov Soyuz I9 15 Jul75 21 Ju175 Leonov AmencadSoviet Test Program Kubasov (ASP). Apollo 18 15 Jul75 24 Jul75 Stafford ASP. Brand Slayton soyuz 20 17 Nov 75 16 Feb 76 Unmanned Soyuz docked to Salyut 4. 22 Jun 76 8 Aug 77 Space station. 6 Jul76 24 Aug 76 Vol ynov Manned Salyut 5. Zholobov 15 Sep 76 23 Sep 76 B ykovsky Independent flight. Aksenov 14 Oct 76 16 Oct 76 Zudov Failed to dock with Salyut 5. Rozhdestensky Manned Spaceflight Summary 215

Designation Launch Recovery Astronauts Notes

Soyuz 24 7 Feb 77 25 Feb 77 Gorbatko Manned Salyut 5. Glazkov Salyut 6 29 Sep 77 Space station. Soyuz 25 9 Oct 77 11 Oct 77 Kovalenok Failed to dock with Salyut 6. Ryumin Soyuz 26 10 Dec 77 16 Jan 78 Romanenko Manned Salyut 6 for 96 days. Grechko Returned in Soyuz 27 sic. Soyuz 2 7 10 Jan 78 16 Feb 78 Dzhanibekov Visited Salyut 6 crew 5 days. Makarov Returned in Soyuz 26 dc. Progress 1 20 Jan 78 8 Feb 78 Progress craft are unmanned supply ships to Soviet space stations. Soyuz 28 2 Mar 78 10 Mar 78 Gubarev Manned Salyut 6. Remek (Czech) Soyuz 29 15 Jun 78 3 Sep 78 Kovalenok Manned Salyut 6 139 days. Ivanchenkov Returned in Soyuz 31 SIC. Soyuz 30 27 Jun 78 5 Jul78 Klimuk Visited Salyut 6 crew. Hermanszewski (Polish) Soyuz 31 26 Aug 78 2 Nov 78 B ykovskiy Returned in Soyuz 29 SIC Jaehn (E. German) Soyuz 32 25 Feb 79 13 Jun 79 Lyakhov Manned Salyut 6 175 days. Ryumin Returned in Soyuz 34 dc. soyuz 33 10 Apr 79 12 Apr 79 Rukavishnikov Failed to dock with Salyut 6. Ivanov (Bulgarian) soyuz 34 6 Jun 79 19 Aug 79 Launched unmanned, returned Salyut 6 crew. SO~UZT-1 16 Dec 79 25 Mar 80 Launched unmanned, docked with Salyut 6. soyuz 35 9 Apr 80 3 Jun 80 Popov Manned Salyut 6 185 days. Ryumin Returned in Soyuz 37 sic. Soyuz 36 26 May 80 30 Jul 80 Kubasov Returned in Soyuz 35 SIC. Farkash (Hungarian) SO~UZT-2 5 Jun 80 24 Jun 80 Malyshev Modified Soyuz spacecraft. Aksenov Docked with Salyut 6.

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Designation Launch Recovery Astronauts Notes

~ ~ Soyuz 37 23 Jul80 11 Oct 80 Gorbatko Returned in Soyuz 36 SIC. Pham (Vietnamese) Soyuz 38 18 Sep 80 26 Sep 80 Romanenko Manned Salyut 6. Tamayo (Cuban) Soyuz T-3 27 Nov 80 10 Dec 80 Kizin 3-man “civilian” Makarov mission to Salyut 6. Strekalov Soyuz T-4 12 Mar 8 1 26 May 8 1 Kovalyenok Manned Salyut 6. Savinykh Soyuz 39 22 Mar 81 Dzhanibekov Manned Salyut 6. Gurragcha (Mongolian) STS I 12 Apr 81 14 Apr 81 Young First Space Shuttle flight. Crippen APPENDIX B United States and Foreign Spacecraft Inventories and Spacecraft Descriptions*

U. S. SPACECRAFT Launch Spacecraft 1st Launch Vehicle Mission AEROS 16 Dec 1972 Scout Upper atmosphere research. AFSATCOM 19 May 1979 Atlas Centaur Air Force Communications (part of FLTSATCOM). Air Density Dec 1963 Scout Explorers 19, 24, 25,39. AMPTE (Active 16 Aug 1984 Delta 3924 Magnetospheric research. Magnetosphere Particle Trace Explorer) ANNA (Army, Navy, 31 Oct 1962 Able Geodetic research (see Secor). NASA, Air Force) satellite Apollo (Manned) 11 Oct 1968 Saturn Lunar exploration, Skylab, Soyuz. AEM-A (Applications 26 Apr 1978 Scout VisuaVLR earth obs. (HCMM). Explorer Mission) AEM-B 18 Feb 1979 Scout Solar radiation study (SAGE).

*Not intended to be a comprehensive listing.

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Launch Spacecraft 1st Launch Vehicle Mission ASC (American 27 Aug 1985 Shuttle Private commercial telmmmunication. Satellite Co.) ATS (Applications 7 Dec 1966 Atlas Agena Geosynchronous experimental Technology Satellites) satellites. (ATS 1 thru 6) Atmosphere 29 Apr 1965 Scout Explorers 27, 32,51,54, 55. Explorer Beacon 23 Oct 1968 Jupiter C Atmospheric research balloons. (Juno) Beacon Explorers 10 Oct 1964 Delta, Scout Explorers S-66,22,27 (geodetic). Big Bird (Broad Jun 1971 Titan 3DI Photohdio reconnaissance. Coverage Photo Agena Reconnaissance Satellites) Biosatellite 15 Feb 1967 Delta Biological research. BMEWS (Ballistic 9 May 1963 Atlas Agena Missile launch detection (MIDAS). Missile Early Warning System) Calsphere 6Oct 1968 Thor Early experimental ASAT. Cameo (Chemically 24 Oct 1978 Delta 2910 Rel. by Nimbus 7 to study . Active Material into Orbit) COBE (Cosmic 1989 Shuttle Big Bang radiation measurement. Background Explorer) COMSAT A company, not a satellite. (Communications Satellite Corporation) Comstar 13 May 1976 Atlas COMSAT telephone communications Centaur satellite. Courier 4Oct 1960 AtlasB Post-SCORE Army communications satellite. CRL (Cambridge 28 Jun 1963 Scout Geophysics research. Research Laboratory) CRRS (Combined 1985 Shuttle Radiation effects measurement, Radiation and Release Satellite) Discoverer 28 Feb 1959 Thor Agena Early reconnaisance satellites. DMSP (Defense Oct 1971 Atlas E Military meteorological satellites. Meteorological Satellite Program) Spacecraft Inventories and Descriptions 219

Launch Spacecraft 1st Launch Vehicle Mission DSCS (Defense 16 Jun 1966 Titan 3C Global strategic military Satellite communications. Communications System) DSP (Defense 5 May 1971 Titan 3C Ballistic missile launch detection. Support Program) Dynamics Aug 1981 Delta Near-earth space environment. Early Bird 6 Apr 1965 Delta Commercial communications (Intelsat 1). Echo 12 Aug 1960 Thor Delta Passive communicationsballoon. ELINT (Electronic Mar 1962 Various Electronic information gathering. Inteligence) satellites (Hitchikers) ERBS (Earth 5 Oct 1984 Shuttle Solar radiation measurement. Radiation Budget Satellite) ERTS (Earth 23 Jul 1972 Delta VisuaYIR earth observation (Landsat). Resources Technology Satellite) ESSA (US. 3 Feb 1966 Delta TOS (TIROS Operational Satellites) Environmental Science (TIROS 1960-65, TOSESSA Services 1966-69, ITOSNOAA 1970-84, Administration) TIROS-N 1978). satellites EUVE (Extreme 1988 Shuttle Ultraviolet telescope Ultraviolet Explorer) Explorer 31 Jan 1958 Jupiter C 1st American satellite in orbit; Series of 61 scientific research satellites from 1958 thru 1975. Ferret (Code 71 1) pre-1973 Thor Agena Electronic reconnaissance precursor to Rhyolite. FLTSATCOM (U.S. 9 Feb 1973 Atlas Centaur Global UHF military Navy Fleet Satellite communications. Communications) Galaxy 28 Jun 1983 Delta Domestic T.V. Galileo 1989 Shuttle Jupiter orbiter. Gemini (manned) 23 Mar 1965 Titan 2 Precursor to Apollo. Geodetic Explorer 6 Nov 1965 Delta Explorers 29, 36 (GEOS A,B) GEOS (Geodynamic 9 Apr 1975 Delta Measurement of earth shape. Experimental Ocean Satellite) (table continued on next page) 220 Introduction to Space Sciences and Spacecraft Applications

Launch Spacecraft 1st Launch Vehicle Mission Geosat 12 Mar 1985 Atlas E Measurement of earth shape. GOES (Geostationary 16 Oct 1975 Delta 2914 Global meteorological information. Operational Environmental Satellite) GPS (Global 22 Feb 1978 Atlas E/F 3-D position and speed fix Positioning System) (NAVSTAR). GRO (Gamma Ray 1988 Shuttle Gamma ray telescope. Observatory) G-STAR May 1985 Ariane3 U.S. commercial domestic communications. HCMM (Heat 26 Apr 1978 Scout D Visual/IR earth observations. Capacity Mapping Mission) HEAO (High Energy 12 Sep 1977 Atlas Centaur Celestial X-ray, Gamma ray map. Astronomy Observatory) 10 Dec 1974 Titan 3E Close solar observation. Hilat 27 Jun 1983 Scout Ionospheric research. HS 376 11 Nov 1982 Shuttle Hughes commercial (or others) communications satellite. (Galaxy, SBS, Telstar, Westar). HST (Hubble 1989 Shuttle Orbiting optical telescope. Space Telescope) IMEWS (Integrated 5 May 1971 Titan 3C Geostationary infrared detection. Missile Early Warning Satellite) Intelsat (International In cooperation with COMSAT Telecommunications Satellite Consortium) Intelsat (International 6 Apr 1965 Delta, Atlas Series of communications satellites Telecommunications Centaur (Blocks 1-6). Satellite Consortium) satellites IUE (International 26 Jan 1978 Delta Celestial UV observation. Ultraviolet Explorer) IMP (Interplanetary 26 Nov 1963 Delta Measure interplanetary environment Monitoring Platform) (Explorers 18, 21,28, 33-35, 41,43,47, SO). Ionds 22 Feb 1978 Nuclear detection satellite. Ionosphere Explorer 25 Aug 1964 Scout Upper atmospheric observation (). Spacecraft Inventories and Descriptions 22l

Launch Spacecraft 1st Launch Vehicle Mission IRAS (Infrared 2.5 Jan 1983 Delta Celestial infrared observation. Astronomical Satellite) ISEE (International 22Oct 1977 Delta Earth-sun space environment. Sun-Earth Explorer) ISIS (International 30 Jan 1969 Delta Ionospheric observation. Satellites for Ionospheric Studies) ITOS (Improved 23 Jan 1970 DeltaN Weather satellites TIROS Operational (TIROS 1960-65, TOSESSA System) 1966-69, ITOSDIOAA 1970-84, TIROS-N 1978). KH-9/11/12 29 Jul 1966 Titan 3D/ Photo reconaissance satellites. (Key Hole) Agena LAGEOS (Laser 4 May 1976 Delta 2914 Passive laser reflector for ranging. GeodynamicsSatellite) Landsat (ERTS 23 Jul 1972 Delta Visual and IR earth observation. Earth Resources Technology Satellite) Leasat 1 Sep 1984 Shuttle Leased FLTSATCOM capability. LDEF (Long Duration 8 Apr 1984 Shuttle Space exposure data. Exposure Facility) LOFT1 (Low 21 Feb 1961 Thor Navigation satellite research. Frequency Trans- (Piggyback) Ionospheric Satellite) Lunar Orbiter 10 Aug 1966 Atlas Agena Lunar topographical data. Lunar Roving Vehicle Moon Buggy. (Magnetic 30 Oct 1979 Scout Near-earth magnetic field data. Field Satellite) Mariner 22 Jul 1962 Atlas Agend Planetary exploration. Centaur 19 Feb 1976 Delta Maritime communications. Mercury 5 May 1961 Redstone/ 1st American manned missions. Atlas Meteoroid Satellites 30 Jun 1961 Scout Explorers S-5.5, 13, 16,23,46. MIDAS (Missile 26 Feb 1960 Atlas Agena IMEWS precursor (BMEWS). Alarm Defense System) MILSTAR 1988 Shuttle Military communications. NATO (NATOSAT) 20 Mar 1970 Delta NATO geostationary communications satellites.

(table continued on next page) 222 Introduction to Space Sciences and Spacecraft Applications

Launch Spacecraft 1st Launch Vehicle Mission Navsat (Naval Operational name of Transit system. Navigation Satellite) NAVSTAR 22 Feb 1978 Atlas E/F 3-D position and speed fixing. (Navigation System Using Time and Ranging)/GPS NERVA (Nuclear Nuclear-powered heavy lift vehicle Engine for Rocket (see RFD) Vehicle Application) Nimbus 28 Aug 1964 Thor Agend Weather satellites. Delta NOAA (National 27 Jun 1979 Atlas ITOS meteorological follow-on Oceanic and (TIROS 1960-65, TOSESSA Atmospheric 1966-69, ITOSMOAA 1970-84, Administration) TIROS-N 1978). satellites NOSS 30 Apr 1976 Naval ocean surveillance system NOVA 15 May 1981 Scout Improved Transit navigational satellite (TIP). NTS (Navigation 14 Jul 1974 Atlas F Tirnatioflransit improvement. Technology Satellites) OAO (Orbiting 8 Apr 1966 Atlas Agend Observation of interstellar space. Astronomical Centaur Observatory) OF0 (Orbiting 9 Nov 1970 Scout Biological experiments on frogs Frog Otolith) in space. OGO (Orbiting 6 Sep 1964 Atlas Agena Study of geophysical and solar Geophysical phenomena. 0bservatory)ROGO OSCAR (Orbiting 23 Jan 1970 Piggyback Ham radio operator broadcasts. Satellite-Carrying Amateur Radio) OS0 (Orbiting Solar 7 Mar 1962 Delta Sun and solar physics observation. Observatory) PAGEOS (Passive 24 Jun 1966 Thor Agena Sunlight reflector for positioning. GEOS) Pegasus 16 Feb 1965 Saturn 1/ Micrometeorite measurements. Apollo Pioneei 11 Oct 1958 Thor, Juno, Lunar, solar, and planetary Atlas, Delta, (Jupiter, Mars, and Venus) probes. Atlas Centaur Probe A/B 19 Oct 1961 Scout Electron density measurement. Project 78 24 Feb 1979 Atlas EartWsun gamma-ray and solar dyn. Spacecraft Inventories and Descriptions 223

Launch Spacecraft 1st Launch Vehicle Mission RMS (Radiation 9 Nov 1970 Scout Meteoroid direction, speed, and flux. Meteoroid Satellite) ,49. Explorer RAM Re-entry 19 Oct 1967 Scout Reentry effects on communication. Experiments Ranger 23 Feb 1961 Atlas Agena Lunar pictures until impact. RCA Satcom 13 Dec 1975 Delta 3914 Commercial communications/TV. Reentry 1 Mar 1962 Scout Reentry effects on ablative material. Relay 13 Dec 1962 Thor Delta Low-altitude repeater . RFD (Re-entry Flight 22 May 1963 Scout Reentry effect on nuclear reactor. Demonstrator) (NERVA) Rhyolite 6 March 1963 Geosynchronous electromagnetic reconnaissance. SAGE (Stratospheric 18 Feb 1979 Scout Solar radiation in earth atmosphere. Aerosol and Gas Experiment) SAMOS (Satellite and 1 I Oct 1960 Atlas, Titan Polar orbiting TV reconnaissance. Missile Observation Agena System) Satcom RCA Satcom. SBS (Satellite 15 Nov 1980 Delta, Shuttle Digital business telecommunication. Business System) SCATHA (Spacecraft 30 Jan 1979 Delta 2914 Geomagnetic field effects on s/c. Charging at High AI titudes) SCORE (Signal 18 Dec 1958 Atlas B Merry Christmas from President Ike. Communication by Orbiting Relay Equipment) Seasat 26 Jun 1978 Atlas F Active/passive oceanographic. Secor (Sequential 11 Jan 1961 Thor Agena Army navigatiodpositioning. Collation of Range spacecraft) SERT (Space Electric 20 Jul 1964 Scout, Thor Ion engine experiments. Rocket Test) Agena Skylab 14 May 1973 Saturn V 1st U. S. space station. Small Astronomy Explorer 53. Satellite

(table continued on next page) 224 Introduction to Space Sciences and Spacecraft Applications

Launch Spacecraft 1st Launch Vehicle Mission Small Scientific . Satellite SME (Solar 6 Oct 1981 Delta Atmospheric research. Mesosphere Explorer) SMM (Solar 14 Feb 1980 Delta 3910 , radiation, particle Maximum Mission) observation. SMS (Synchronous 17 May 1974 Delta First GOES satellite. Meteorological Satellite) Solar Explorer Explorer 30, 37. SOLRAD Explorer 44. Spacenet 23 May 1984 Ariane U.S. commercial domestic communications. STC (Satellite 1987 Ariane U.S. television broadcasts. Television Corp.) STS (Space Shuttle) 12 Apr 1981 SRBs Space Transportation System. Squanto Terror Mar 1964 Thor Once-operational nuclear ASAT. (Johnson Is.) Surveyor 30 May 1966 Atlas Soft lunar lander. CentaurIAgena SMS (Synchronous 17 May 1974 Delta Precursors of GOES, Meteorologicd Satellites) Syncom 14 Feb 1963 Thor Delta 1st geosynchronous communications satellites. Tacsat 9 Feb 1969 Titan 3C Military tactical communications. TDRS (Tracking and 4 Apr 1983 ShuttlelIuS SIC tracking and data relay. Data Relay Satellite) Teal Ruby 1987 IR aircraft detection system. Telstar 10 Jul 1962 Thor Delta 1st commercial communications satellite. Timation (Time May 1967 scout Precursor to GPS. Navigation) TIP (Transit Impruved 2 Sep 1972 scout Navigational satellite with Pr0gram)iNOVA DISCOSRTGs. TIROS (Television 1 Apr 1960 Thor, Delta, Infrared weather monitoring Infrared Observation Atlas (TIROS 1960-65, TOSlESSA Satellite) 1966-69, ITOSNOAA 1970-84, TIROS-N 1978). TOP0 8 Aur 1970 Thor Agena Army space-ground triangulation. Spacecraft Inventories and Descriptions 225

Launch Spacecraft 1 st Launch Vehicle Mission TOS (TIROS 3 Feb 1966 Delta ESSA meteorological satellites Operational System) (TIROS 1960-65, TOSESSA 1966-69, ITOSMOAA 1970-84, TIROS-N 1978). TRAAC (Transit 15 Nov 1961 Thor Gravity gradient research. Research and Attitude (Piggyback) Control) satellite Transit Sep 1959 Scout Navigational satellite system. UARS (Upper 1989 Shuttle Stratosphere, mesosphere monitor. Atmosphere Research Satellite) Vanguard 6 Dec 1963 Vanguard Early Navy satellite program. Tv-3 Vela 17 Oct 1963 Atlas Agena Nuclear explosion detection. Viking 20 Aug 1975 Titan Centaur Soft landing on Mars. Voyager 20 Aug 1977 Titan Centaur Planetary exploration. Westar 13 Apr 1974 Delta 2914 I st domestic geosynchronous communications satellites. Whitecloud 30 Apr 1976 Atlas Electronic ocean surveillance/ tracking. NON-11. S. SATELLITES

~~ ~ Launch Spacecraft 1st Launch Vehicle Mission

Aeros 16 Dec 1972 Scout W. German environment tests. Alouette 29 Sep 1962 Thor Agena Canadian ionospheric effects on communications Anik-B 15 Dec 1978 Delta Canadian telecommunications. ANS (Astronomical 30 Aug 1974 Scout Netherland X-ray studies. Netherlands Satellite) Apple (Ariane 19 Jun 1981 Ariane Indian communications experiments. Passenger Payload Experiment) Arabsat 8 Feb 1985 Ariane, Arab League telecommunications. Shuttle Ariel 27 Mar 1964 Scout U. K. radio and environment tests. Aryabhata I9 Apr 1975 C-1 Indian high-altitude physics. Astron 23 Mar 1983 D-1 Soviet X-rayKJV telescope. Aussat 27 Aug 1985 Shuttle Australian telecommunications.

(table continued on next page) 226 Introduction to Space Sciences and Spacecraft Applications

Launch Spacecraft 1 st Launch Vehicle Mission

AyameECSs 6 Feb 1979 N- 1 Japanese communications. (Experimental Communications Satellite) AZUR 8 Nov 1969 scout W. German particle measurements. Bhaskara 7 Jun 1979 C-1 (Sov) Indian Earth resources satellite. Brasilsat 8 Feb 1985 Ariane Brazilian telecommunications. China- 1 24 Apr 1970 Long March- 1 Chinese space environment tests. CORSA (Cosmic 21 Feb 1979 M-3C-4 Japanese cosmic radiation Radiation Satellite) experiment. COS-B 9 Aug 1975 Delta ESA gamma ray observations. COSMOS- 1076 12 Feb 1979 Soviet oceanographic satellite. (Most Soviet spacecraft are given the generic label COSMOS until typed). Cosmos 1629/1700 21 Feb 1985 Soviet TDRS-type relay satellite. Cosmos-Meteor 25 Jun 1966 A-2 Soviet weather satellite. Cosmos-Tsikada 23 Nov 1967 C- 1 Soviet Transit-type navigational satellite. CTS (Communications 17 Jan 1976 Delta 2914 USA/Canada/ESA communications Technology Satellite)/ satellite. Hermes DENPNREXS 19 Aug 1972 M-4s-4 Japanese magnetic field experiment. DIALNIKA 10 Mar 1970 B W. German particle measurements. DiapasodDiademe 17 Feb 1966 Diamant French geodetic satellites. ECS (European 16 June 1983 Ariane European continental Communications communications satellite. Satellite) EGP (Experimental 8 Aug 1986 H- 1 Japanese geodetic research. Geodetic Payload) Ekran 26 Oct 1976 D-1 Soviet TV broadcast. Elektron 30 Jan 1964 A- 1 Soviet radiation belt experiments. EOLE FR-2 16 Aug 1971 scout French atmospheric data gathering. (Vandenberg) ERS (European 1989 Ariane European remote sensing satellite. Remote Sensing) satellite ESRO (European 17 May 1968 ESA space physics experiments. Space Research Organization) Spacecraft Inventories and Descriptions 227

Launch Spacecraft 1st Launch Vehicle Mission

ETSlKiku 23 Feb 1977 N-1 1 st Japanese geostationary test (Engineering Test satellite. Satellite) Exos 4 Feb 1978 M-3H-2 Japanese particle measurements. EXOSAT (European 26 May 1983 Delta 3914 ESA X-ray observations. X-ray Observatory Satellite) FR- 1 6 Dec 1965 Scout French ionosphere/magnetosphere. GEOS 20 Apr 1977 Delta ESA magnetospheric studies. GLONASS (Global 12Oct 1982 D-1 Soviet GPS-type navigation system. Navigation Satellite System) GMS/Himawari 14 Jul 1977 Delta Japanese meteorological satellite. (Geostationary Meteorological Satellite) Gorizont 19 Dec 1978 D-1 Soviet geostationary communications/TV satellite. HEOS (Highly 5 Dec 1968 Delta ESA solar/magnetosphere Eccentric Orbit experiment. Satellite)

Hinoton/ Astro 21 Feb 1981 M-3.3-2 Japanese solar flare observations. INSAT-B (Indian 30 Aug 1983 Shuttle Indian meteorological/ National Satellite) communications satellite. Intasat 15 Nov 1974 Delta Spanish signal propagation experiment. ISEE-2 (International 22 Oct 1977 Delta ESA portion of ISEE experiment. Sun-Earth Explorer) ISIS (International 30 Jan 1969 Thor Delta Alouette follow-on. Satellites for Ionospheric Studies) Italsat I989 Ariane Italian telecommunications. JERS (Japanese 1991 H- 1 Japanese SAR remote sensor. Earth Resources Satellite) Marecs 20 Dec 1981 Ariane ESA maritime communications. Meteor- 1 26M~1969 A-2 Soviet weather satellite. Meteor-2 11 Jul 1975 A-1 Soviet meteorological- satellite.

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Launch Spacecraft 1 st Launch Vehicle Mission

Meteor-Priroda 10 Jul 1981 A-1 Soviet earth resources satellite. Meteosat 23 Nov 1977 Delta 2914 ESA geosynchronous meteorological satellite. Molniya 23 Apr 1965 A-2 Soviet telecommunications. Morelos 17 Jun 1985 Shuttle Mexican telecommunications. MOS (Marine 1987 N-2 Japanese Earth observation satellite. Observations Satellite) Olympus (L-Sat) 1988 ESA advanced telecommunications experiment. OREOUAUREOLE 27 Dec 1971 C-l SovietRrench solar wind experiment. OTS (Orbital Test 12 May 1978 Delta 3914 ESA experimental communications Satellite) satellite. Palapa 8 Jul 1976 Delta Indonesian telecommunications. 18 Jun 1983 Shuttle Peole 12 Dec 1970 Diamant B French geodetic/transmission tests. PolluxKastor 17 May 1975 Diamant French satellite technology B-P4 experiment. Prognoz 14 Apr 1972 A-2 Soviet solar studies. 16Jul 1965 D Soviet high-energy particle experiment. Radarsat 1991 Shuttle Canadian SAR remote sensor. Radio 26 Oct 1978 F-2 Soviet ham radio satellite. (Piggyback) Raduga 22 Dec 1975 D-1 Soviet geosynchronous communications satellite. SakurdCS 15 Dec 1977 Delta Japanese experimental (Communications communications satellite. Satellite) San Marco 15 Dec 1967 Scout Italian atmospheric research satellites. Shinsei 28 Sep 1971 M-4s-3 Japanese cosmic ray measurement. Signe 17 Jun 1977 C-1 French X-ray, gamma ray experiment. Sirio 25 Aug 1977 Delta Italian experimental communications satellite. SPAS (Shuttle 18 Jun 1983 Shuttle W. German reusable experiment Pallet Satellite) platform. SPOT (Satellite 22 Feb 1986 Ariane-1 European earth imaging system. Probatoire de I’Observation de la Terre) Spacecraft Inventories and Descriptions 229

Launch Spacecraft 1st Launch Vehicle Mission

Sputnik 4 Oct 1957 Type A 1 st artificial satellite (Soviet). SRATS (Solar 24 Feb 1975 M-3C-2 Japanese solar radiation experiment. Radiation and Thermospheric Structure) satellite Starlette 6 Feb 1975 Diamant French geodetic satellite. B-P4 STW 8 Apr 1984 CZ-3 1st geostationary Chinese communications satellite. Symphonie 19 Dec 1974 Thor Delta FrencNW. German telecommunications satellite. TDFRV-SAT 1987 Ariane FrencNW. German direct TV. Telesat 4 Aug 1984 Ariane French telecommunications. Tele-X 1988 Ariane Scandinavian direct TV/data. TenmdAstro 20 Feb 1983 M-3s-3 Japanese X-ray observations. TOPEX (The Ocean 1991 Ariane U.S./French ocean altimetry. Topography Experiment) satellite Tournesol/Aura 15 Apr 1971 Diamant B French solar radiation experiments. UME/ISS (Ionospheric 29 Feb 1976 N-1 Japanese ionospheric experiments. Sounding Satellite) Viking 22 Feb 1986 Ariane Swedish space plasma experiments. YudBSE 8 Apr 1978 Delta Japanese direct TV broadcast. (Broadcasting Satellite Experiment) INDEX

3-axis stabilization, 171 Aristotle, 26 Absolute time reference, 162 Armstrong, Neil, 7 Absorption bands, 145 Astronomical unit, 8 1 Absorptivity, 183 Atlas, 170 Acceleration, 56 Atmospheric composition, 89 Active sensor, 133 Atmospheric density, 91,184, 187 Adiabatic flame temperature, 59 Atmospheric drag, 185, 187 Aerosol, 143 Atmospheric temperature, 90 Airborne launches, 193 Atmospheric windows, 72, 145, 190 Alternative design solutions, 202 Attitude reference and control Altimeter, 136 (ARC), 172 Altitude error, 158 Backscatter, 142 Amplitude, 105 Baikonur cosmodrome, 72 modulation (AM), 106, 116 Ballistic cross-section, 169, 187 shift keying (ASK), 126 Ballistic Missile Defense Organization Analog signals, 115 (BMDO), 22 Analog to digital conversion, 120 Bandwidth, 114 Angular field-of-view, 137, 139 Baseband signals, 115 Angular momentum, 34 digital, 123 Anisotropic radiator, 109 Battery type, 179 Announcement of opportunity Beginning of life power (BOL), 179 (AO), 205 Biphase modulation, 160 Antenna, 108, 188 Bipropellent, 185 effective area, 111 Binary, 119 efficiency, 11 1 Bit, 120 gain, 110 Bit period, 123 patterns, 109 Blackbody, 80 Aperture synthesis, 148 Body axes, 174 Apoapsis, 29, 31 Body-stabilized, 174 Apogee, 33 Boresight, 137 kick motor (AKM), 72 Brahe, Tycho, 26 Apollo, 7 Bum-out, 64 Apollo-Soyuz, 9 Byte, 120 Applied Physics Laboratory (APL), 154,193 Camera, 133 Argument of perigee, 38 Carrier wave, 115 Ariane. 22 Centaur, 194

230 Index 231

Charge coupled devices (CCDs), 136 mass iteration, 191 Chemical power, 177 process, 199, 201 Chromosphere, 79 reviews, 204 Circular orbit, 30 Diffraction limit, 138 Clarke, Arthur, 2, 104 Digital baseband signal, 160 Clock offset time, 162 Digital signals, 115 Closed orbits, 28 Direct waves, 106 COBE, 18 DMSP, 14, 134 Command processing, 190 Doppler Commands, 189 count, 155 Communication relay satellite, 188 frequency, 155 Complex autonomy, 190 positioning, 154 Comsat, 19 shift, 154, 158 Conceptual study phase Drag, 100 (Pre-Phase A), 200, 204 Dual-spin, 172 Conical scan, 140 Duty cycle, 179, 189 Constellations, 46 Dwell time, 148 Control torques, 176 Dynamic mass estimate, 169 Convective zone, 78 Copernicus, 26 Earth, 87 Earth angle, 137 Corona, 79 Earth Observing System (EOS), 23 Correlation, 161 Corrosion, 98 Eastern test range (ETR), 193 Cosmology, 76 Eccentricity, 29 cost, 201 Eclipse cycle, 179 Eclipse periods, 169 Critical design review (CDR), 207 Ecliptic, 77 Cross correlation, 161 Cross-track scan, 140 Electromagnetic propagation, 91, 105 Electromagnetic spectrum, 105 Data handling, 188 Electromagnetic wave, 104 Data rate, 189 Electron density, 107 Dead reckoning (DR), 157 Electrostatic system, 185 Debris, 99 Elevation angle, 137 Decibel, 109 Elliptical orbit, 3 1 Definition phase (Phase B), 206 ELV, 21 Delta, 170 Emissivity, 183 Delta v, 48, 70, 71, 73, 186 End of life power (EOL), 179 Demodulation, 118, 128 , 21 Depression angle, 137 Ephemeris, 155 Depth of discharge, 181 Equitorial orbit, 44 Design Equivalent noise temperature, 113 acquisition period, 200 ERTS, 14 development phase (CAI), 207 European Space Agency (ESA), 21 232 Introduction to Space Sciences and Spacecraft Applications

EVA, 6 Grazing angle, 138 Exhaust velocity, 57,58 Greenhouse effect, 92 Exosphere, 91 GRO, 18 Explorer, 15 Ground instantaneous field of view Explorer 1, 2, 54 (GIFOV), 139 Explorer, 15, 132 Ground sites, 194 Field-of-view (FOV), 39,40, 137, 146 Ground support systems, 192 Final orbit, 186 Ground track, 41 Footprint, 40 Ground waves, 106 Free space loss, 113 Guiana launch site, 72 Frequency, 79 Gyroscopic stabilization, 174 modulation, 106, 116 HEAO, 18 shift keying, 127 Heliocentric, 26 Fuel mass, 187 Heliosphere, 85 Gagarin, Yuri, 5 Hohmann transfer, 49 Galileo, 26, 27 Horizon sensors, 172 Galileo, 17, 18 HST, 18 Gemini, 5,6 Hydrazine, 74, 185 GEO, 185 Hydrazinehitrogen tetroxide, 185 Geodesy, 15 Hyperbolic orbit, 30 Geoid, 158 ICBM, 3,20 Geomagnetic field, 93 Imager, 135 Geomagnetic storms, 95 Incidence angle, 138 Geophysical processes, 135 Inclination, 38, 67, 69, 72 Geostationary orbit, 19,44 Index of refraction, 142 Geostationary transfer, 7 1 Inertial Upper Stage (IUS), 194 Geosynchronous orbit (GEO), 43, 172 Information bandwidth, 189 Geosynchronous satellites, 187 Infrared, 133 Global Positioning Satellite (GPS), Instantaneous field-of-view (IFOV), 22, 154, 163 137, 146 Goddard Space Flight Center (GSFC), Intelsat, 19 193 International space station, 22, 100 Goddard, Robert, 3,53 Ionosphere, 87 GOES, 14,15, 134 Ionospheric effect on electromagnetic Gorizont, 20 propagation, 106, 107 Granulation, 78 ISEE, 15 Gravitational acceleration, 187 Isotropic radiator, 109 Gravitational field, 185 IUE, 18 Gravitational parameter, 32 Gravity gradient stabilized, 171 Jet Propulsion Laboratory (JPL), 193 Gravity gradient boom, 172 Johns Hopkins University, 154 Index 233

K-band, 188 Mesosphere, 90 Kennedy Space Center (KSC), 69, Meterological satellites, 132 193 Micrometeorites, 99 Kepler, 26 Microwave, 133 Kepler’s laws, 28 bands, 105 MILSTAR, 22 Landsat, 14, 133 Mir; 12, 13, 22 Laser, 133 Mission analysis phase, 200 Launch, 192 Mission from Planet Earth, 23 azimuth, 67 Mission lifetime, 169, 187 configuration, 191 Mission need, 199 sites, 73, 193 Mission to Planet Earth, 23 vehicles, 170, 193 Momentum, 55 weight, 193 Momentum wheels, 172 window, 66 Multiple-use designs, 195 LDEF, 16 Multistage rockets, 194 LEM, 7 Lift-off thrust, 193 Nadir, 39 Light-year, 77 angle, 137 LIGHTSATs, 195 National Academy of Science, 199 Link budget, 111 National Advisory Committee for Liquid-fueled propulsion, 186 Aeronautics (NACA), 4 Liquid oxygen (LOX), 59 National Aeronautics and Space Longitude of the ascending node, 38 Administration (NASA), 4 Look angle, 137 National Aerospace Plane (NASP), 21 Loran, 163 National Oceanic and Atmospheric Low-earth orbit (LEO), 44 Administration (NOAA), Lunar Orbiter, 17 14, 134 Magellan, 17 National Space Council, 199 Magnetic tape storage device, 189 Natural radiation, 135 Magnetic torquers, 176 Navigation, 2 Magnetopause, 94 control, 190 Magnetosphere, 95 message, 155, 162 Man in space, 101 NavstadGPS, 21 Mariner, 17 Newton, 27 Mars Observer; 74 Newton’s laws, 34 Mass dynamics, 166 Noise, 113 Mass flow rate, 57 Northern lights, 96 Mass imbalance, 191 Nozzles, 61 Mass ratio, 65 Nuclear power, 177 Mercury, 5 Nuclear reactor, 177 Mesopause, 91 Nybble, 120 234 Introduction to Space Sciences and Spacecraft Applications

OAO, 18 Periapsis, 29, 31 Oberth, Hermann, 3,54 Perigee, 33 Omega, 163 Perigee kick motor (PKM), 7 1 Omnidirectional antennas, 189 Phase, 80, 105 On-orbit mass, 187, 188 Phase difference positioning, 159 Onboard computer subsystem, 190 Phase modulation, 117 Onboard recorders, 188 Phase shift keying, 127 Onboard recording, 189 Photodissociation, 89 Open orbits, 30 Photoionization, 88 Operational configuration, 191 Photosphere, 77,79 Operations phase (Phase E), 209 Picture element (pixel), 138 Optimization, 203 Pioneer, 16 Orbit Pitch, roll, and yaw, 174 adjustment, 70 Planck’s law, 82 establishment, 47 Plane change, 71 maintenance, 186,190 Planning period, 200 properties, 39 Pointing accuracy, 169, 172 Orbital Polar orbit, 45 altitude, 37 Polarization, 142 elements, 38 Power maneuvers, 48 budget, 111 parameters, 28 control unit, 181 period, 32 management, 190 radius, 30 profiles, 166 translation, 186 regulation, 181 velocity, 36 Preliminary design review (PDR), 207 OSO, 16 Pressure thrust, 62 Outgassing, 99 Pro-grade orbit, 68 Overexpansion, 62 Programmatic, 204 Ozone depletion, 134 Proof-of-concept, 134 Propagation velocity, 142 Parabolic dish antenna, 110 Propellant mass flow rate, 57 Parabolic orbit, 30 Propellants, 60, 185 Parking orbit, 186 Pulse length, 146 Passive sensor, 133 Pulse ranging, 146, 154, 159 Payload, 166 Pushbroom scan, 140 management, 190 specialist, 205 Q-levels, 121 support dynamics, 191 Quality assurance, 204 Payload Assist Module (PAM), 194 Quantization, 121 Pegasus XL, 171 Radar, 133 Perfect expansion, 63 Radians, 146 Index 235

Radiation belts, 96 Sampling, 122 Radiation effects on spacecraft, 98 frequency, 122 Radiative transfer equation, 142, 145 Satellite Radiative zone, 78 design, 166 Radio bands, 105 subsystems, 166, 167, 192 Radioactive, 177 vital statistics, 169 Radioisotope thermal generators Saturn V, 7 (RTGs), 177 Scanning payloads, 19 1 Radiometer, 136 Scatterometer, 136 Radiometric, 136 SDI, 22 Range, 146 Seasat, 16, 133 cross-track, 139 Semi-major axis, 29 Ranger, 17 Sensor scanning, 140 Ranging signal, 160 Sensors, 100 Ratio of specific heats, 59 Shepard, Alan, 5 Rayleigh scattering, 144 Sidereal day, 32 Reaction control wheels, 172 Signal propagation velocity, 146 Real aperture radar (RAR), 136, 147 Signal-to-noise ratio, 1 14 Receiving sites, 194 Simple autonomy, 190 Recombination, 177 Sky waves, 106 Reflectance, 149 Skylab, 7, 100 Refraction, 106 Slant range, 137, 155 errors, 157, 163 SMALLSATs, 195 Refrigeration devices, 184 Solar Remote sensing, 1 cells, 177 Replica code, 161 constant, 81 Request for proposal (RFP), 193 cycle, 86 Resolution cell, 138 cycle activity, 184 Retrograde orbit, 68 day, 32 Revisit time, 43 eclipse times, 180 Revolutions per day, 42 electromagnetic radiation, 79 Risk assessment, 205 flares, 86 Rocket equation, 57 heat exchangers, 177 Rocket thrust, 57 maximum, 86 Rotisserie, 183 particle radiation, 84 Rough order of magnitude (ROM), power, 177 205 pressure drag, 159 RP- 1. 60 radiation, 79 wind, 84 S-band, 188 SolarMax (SMM), 16 Salyut, 12 Solid-state storage device, 189 Sample period, 122 Solid-fueled propulsion, 186 236 Introduction to Space Sciences and Spacecraft Applications

Sounder, 135 Sunspots, 85 South Atlantic Anomaly, 97 Surface phenomena, 142 soyuz, 12 Surveyor, 17 Space Shuttle, 9, 170, 194 Swath width, 40 Space waves, 108 Synchronization, 161 Spacecraft Synthetic aperture radar (SAR), 15, batteries, 181 17,136, 147 charging, 98 System definition phase, 200 contol facilities, 194 System life cycle, 199 dry weight, 188 System life cycle phases, 200, 203 evolution, 195 System user, 194 performance, 100 Systems approach, 198 systems, 198 Systems engineering, 200 wet weight, 188 SPAWAR, 193 TACAN, 163 Specific energy density, 181 Tankage, 185 Specific gas constant, 59 Target incident angle, 147 Specific impulse, 60, 185, 187 TDRSS, 188 Spectrometer, 136 Telemetry processing, 190 Spheroid surface, 158 Telemetry signal, 188 Spin-stabilized, 171 Temporal, 149 SPOT, 19 Thermal control, 97 Sputnik 1,2, 3,54 Thermal gradients, 171 SSME, 61 Thermal subsystem, 183 SSTO, 21 Thermosphere, 91 Stable frequency, 155 Thrust, 57 Staging, 65 Thrusters, 172, 185 Star mapper, 176 Time, 201 Star trackers, 172 Time-in-view, 42 Static mass estimate, 169 Time/frequency, 136 Station-keeping, 72, 186 Timing signals, 155 Status monitoring, 192 TIROS, 14 Stefan-Boltzmann, 8 1 Tiros 1, 14, 132 Stratopause, 90 Titan, 170 Stratosphere, 90 TOMS, 14 Structure subsystem, 190 TOPEXPoseidon, 15 STS, 9 Total energy, 35 Subsystem mass values, 192 Tracking, 192 Sun (see Solar) Trade studies, 205 Sun composition, 77 Trade-offs, 201 Sun sensors, 172 Transfer orbit, 7 I Sun-synchronous orbit, 45 Transit, 154, 156, 163 Index

Transmission rate, 122 Van Allen, James, 3,96 Transmissivity, 144 Vandenberg launch site, 69, 193 Transponder, 188 Velocity vector, 186 Tropopause, 90 Venera, 18 Troposphere, 90 Viking, 17, 194 Tropospheric errors, 158 Visible, 133 True anomaly, 30 Volume phenomena, 142 Tsiolkovsky, Konstantin, 3,53 von Braun, Wernher, 3,54 TT&C, 188 Voskhod, 11 Quratam launch site, 72 Vostok, 5,ll Voyager, 16, 194 UARS, 14 Ultraviolet (UV), 149 Wavelength, 79 ULysses, 16 Weather information, 194 Unbalancing torques, 172 Weather satellites, 132 Uncompensated momentum, 191 Western test range (WTR), 193 Underexpansion, 62 Wien’s law, 82 Universal Gas Constant, 59 Wiring mass, 190 Upper stage, 194 Use period, 200 X-band, 188 v-2, 3, 20, 54 Zenith angle, 142 Vacuum, 98 Almanac of I Soviet Manned Space Flight Dennis Newkirk I “...wall-to-wall Here is a complete and up-to-date account information, all of every Soviet manned space launch. of it quite 378 pages, photos, illustrations, appendix, index, I fascinating...” 6” x 9” hardcover. -Popular $29.95 ISBN 0-87201-848-2 #1848 I Communications Dead Reckoning: I Calculating Without Instruments Ronald M! Doerfler “...reasoned set of stratagems for fast mental I Space Satellite Handbook calculations ...that allow impressive numerical I Third Edition approximation...” I Anthony R. Curtis ScientificAmerican 182 pages, figures,tables, appendix, references, The third edition of this unique encyclope- 1 index, 6” x 9” paperback. I dia describes more than 22,000 man-made $17.95 ISBN 0-88415-087-9 I objects launched into orbit from all coun- #5087 I tries since 1958. You get the latest data from SOLVE IT! NASA and other agencies including each Computer-Aided Mathematics for 1 I satellite’s official international number, popular name, launch date, and country of Science and Engineering Samuel Doughty, Ph.D., RE. I origin. I 358 pages, figures, charts, diagrams, tables. This book shows readers how to develop 1 maps, timetables, calendar, index. 8‘/P x 11” their own ability to write mathematical 1 I ;arkover. computer programs. ISBN 0-88415-192-1 #5192 164 pages, figures, index, 8%” x lo7/: hardcover. $29.95 ISBN 0-88415-266-9 #5266 I Space Almanac 1 Second Edition Visit Your Favorite Bookstore! I Anthony R. Curtis Or order directly from: 1 “...packed with more fun and interesting facts Gulf Publishing Company I about space exploration....” C P.O. Box 2608 Dept. IK Sky& Telescope Houston, Texas 77252-2608 1 746 pages, index with 4,000 entries, figures, 0 713-520-4444 FAX: 713-525-4647 I charts, diagrams, tables, maps, timetables. E-Mail: [email protected] I calendar, 6 x 9’: Send payment plus $4.95 ($6.15 if the order is $20 or I $24.95 paperback more, $7.75 for orders of $30 or more) shipping and han- ISBN 0-88415-030-5 #5030 dling or credit card information. CA, IL, NJ, PA, and TX I $36.95 hardcover residents must add sales tax on books and shipping total. I I ISBN 0-88415-039-9 #5039 Thank you for your order! I